University Student Launch Initiative 2017-2018 Critical ... · Justification for material...

NORTHEASTERN

UNIVERSITY University Student Launch Initiative

2017-2018

Critical Design Review

January 12, 2018

Northeastern University 2017-2018 Student Launch Critical Design Review 1

Table of Contents 1. Team Summary ........................................................................................................................... 3

1.1. Team Summary ................................................................................................................ 3

1.2 Launch Vehicle Summary ................................................................................................ 3

1.3 Payload Summary ............................................................................................................ 3

1.4. Changes Made Since PDR ................................................................................................... 4

1.4.1. Changes Made to Vehicle Criteria ................................................................................. 4

1.4.2. Changes Made to Payload Criteria ................................................................................ 4

1.4.3. Changes Made to Project Plan ....................................................................................... 4

2. Vehicle Criteria ........................................................................................................................... 5

2.1. Design and Verification of Launch Vehicle ......................................................................... 5

2.1.1. Flight Reliability and Confidence .................................................................................. 5

2.1.2 Design Alternatives from PDR ....................................................................................... 5

2.1.3 Computer Aided Design (CAD) Drawings ..................................................................... 7

2.1.4. Discuss the Integrity of Design ...................................................................................... 9

2.1.3. Justification for material selection, dimensioning, component placement, and other

unique design aspects ............................................................................................................. 12

2.2. Subscale Flight Results ...................................................................................................... 14

2.2.1. Subscale Flight Results ................................................................................................ 14

2.2.2. Perform an Analysis of the Subscale Flight ................................................................. 18

2.2.3. Impact of Subscale Flight Data on Full Scale Launch Vehicle ................................... 22

2.3. Recovery Subsystem .......................................................................................................... 23

2.4. Mission Performance Predictions ....................................................................................... 30

3. Safety ........................................................................................................................................ 36

3.1. Launch Concerns and Operation Procedures ..................................................................... 36

3.2. Safety and Environment (Vehicle and Payload) ................................................................ 40

Likelihood Definitions ........................................................................................................... 54

Severity Definitions ............................................................................................................... 54

Environmental Concerns ........................................................................................................ 55

Table 3.2.5 : Environmental Concerns ...................................................................................... 55

4. Payload Criteria ........................................................................................................................ 57

4.1. Design of Payload Equipment ............................................................................................ 57

4.1.1. Wheels.......................................................................................................................... 57

4.1.2. Ejection ........................................................................................................................ 58

4.1.3. Solar Panel Deployment .............................................................................................. 59

4.2. System Level Design Review ............................................................................................. 59

4.2.1. Electronics.................................................................................................................... 59

4.2.2. Solar Panel ................................................................................................................... 60

4.2.2. Solar Panel ................................................................................................................... 60

4.2.3. Wheels.......................................................................................................................... 60

4.2.3. Payload Ejection System.............................................................................................. 60

4.3. Drawings and Specifications for Components and Assembly ........................................... 60

4.4. Payload Component Interaction Description ..................................................................... 64


4.4.1. Payload Component Interaction Overview .................................................................. 64

4.4.2. NUFR Component Interaction ..................................................................................... 64

4.4.3. PES Component Interaction ......................................................................................... 67

4.5. Payload Integration Plan .................................................................................................... 67

4.6. Demonstration of Payload Design Completion .................................................................. 68

4.7. Payload Electrical Drawings and Dimensions ................................................................... 70

4.8. Payload Block Diagrams .................................................................................................... 72

4.9. Payload Battery and Power Consumption .......................................................................... 74

4.10. Switch and Indicator Wattage and Location .................................................................... 74

4.11. Payload Justification ......................................................................................................... 76

5. Project Plan ............................................................................................................................... 78

5.1. Testing ................................................................................................................................ 78

5.1.1. Payload Tests ............................................................................................................... 78

5.1.2. Rover Test Plans .......................................................................................................... 78

5.2. Requirements Compliance ................................................................................................. 80

5.3. Budgeting and Timeline ..................................................................................................... 92


1. Team Summary

1.1. Team Summary Team name: NU Frontiers Mailing Address: Northeastern University, 267 Snell Engineering, Boston, MA 02115 Mentor: Robert DeHate Certification Level: L3 Nar/TRA #75198/TAP 9956

1.2 Launch Vehicle Summary The Launch Vehicle has been designed to propel itself and the payload secured inside to an apogee of

5,303 feet. The Launch Vehicle will be made of 4 independent sections (Nose Cone, Payload, Lower

Avionics Bay, and Booster) that will separate during flight events controlled from two avionics bays.

Both avionics bays will contain StratoLogger CF altimeters while the lower avionics bay also will also

contain an XBEE Pro XSC (S3) GPS. Power to the avionics bays will be supplied by 9V batteries, with

each altimeter using a dedicated battery. At apogee, the Lower Avionics Bay and the Booster section will

separate from each other while remaining tethered, deploying a 48 inch drogue parachute. At an altitude

of 800 feet, two events will separate the Launch Vehicle into four independent sections. The Nose Cone

section and the Payload section will be tethered together and will descend on a 72 inch main elliptical

parachute. The Lower Avionics Bay and the Booster will be tethered together and will descend using the

previously deployed drogue parachute and another 72 inch elliptical main parachute. The length of the

Launch Vehicle is 148 inches with an outer diameter of 6.17 inches. The Launch Vehicle, including the

loaded motor, will have a mass of 47.4 pounds, and will be propelled by a Cesaroni L1115 Classic

reloadable motor. The Launch Vehicle will be launching off of a 12 foot 1515 rail.

1.3 Payload Summary NU-FRONTIERS The design experiment chosen was Option 2, the deployable rover. A custom rover and deployment

mechanism has been designed to interface with the launch vehicle. The goal of the experiment is to have

the payload secured in the launch vehicle payload bay throughout the flight and remain safely

encapsulated until the end of the flight path. Once the payload section is safely and successfully on the

ground, the rover deployment method will be activated. The deployment method will utilize a custom

pneumatic piston that will guide the rover along linear bearings until the rover has been fully released

from the launch vehicle body tube. Once the rover is fully deployed, the rover onboard electronics will

measure and record position from the starting point and chassis orientation. The motors will drive the two

main wheels and one supporting wheel to a distance of 5 feet from the landing location, overcoming any

obstacles in the terrain. Onboard electronics will sense if there is an obstacle obstruction the rover

trajectory and navigate a new path. When the rover reaches 5 feet, the custom program logic will

autonomously stop the driving motion. At this point, a servo motor will initiate the foldable solar panel

deployment by means of the rotating fan method. The solar energy collectors will start in a folded and

enclosed position within the rover casing and will be expanded to have panels from 0 to 180 degrees for


an increase in surface area. The solar panels will be connected to the onboard electronics bay and when

power is generated an indicator light will blink.

1.4. Changes Made Since PDR

1.4.1. Changes Made to Vehicle Criteria

The most significant changes to the Launch Vehicle between the Preliminary Design Review and

the Critical Design Review include revised mass estimates, a relocation of the payload in the

rocket, changes in the recovery system, and a different motor selection. The Launch Vehicle will

still consist of four sections: the Nose Cone, Payload, Lower Avionics Bay, and Booster sections.

At PDR, the Nose Cone section was planned to separate at 800 feet from the remainder of the

Launch Vehicle and fall independently. This has been changed as the Payload section will now

also separate with the Nose Cone section at 800 feet. This allows the payload to be deployed on

the downwards facing end of section instead of the upwards facing end. The final concept of operations is that, at apogee, the drogue parachute will be deployed through

separation of the Booster and Lower Avionics Bay sections. At 800 feet, two events will occur.

The first event will separate the Nose Cone and Payload sections, deploying a 72 inch main

elliptical parachute that tethers these two sections together. Very shortly after, a second event will

separate the payload and lower avionics sections, releasing another 72 inch main elliptical

parachute tethered to the Lower Avionics Bay and Booster sections. In order to allow the launch vehicle to be assembled more simply, the airframe of the Lower

Avionics Bay section splits in two. The Lower Avionics Bay will now be housed in a 12 inch

section of coupler and will be set screwed in the center of the two split airframe pieces. This will

make the installation of the lower avionics bay much easier as the assembler does not have to

reach their arm down the airframe. With the addition of another coupler tube and a heavier

payload estimate, mass estimation of the Launch Vehicle has increased since PDR leading to a

change in motor selection. The new selected motor is the Cesaroni L1115 Classic reloadable

motor, bringing the Launch Vehicle to a new estimated 5,302 feet. The final launch rail selection

will be a twelve foot 1515 rail. 1.4.2. Changes Made to Payload Criteria

No major design changes were made to the payload since PDR. The design has been refined and

the specifics such as which motors, servos, and electronics would be used were figured out. The

initial concept of a two wheeled rover with a folding counter torque tail was maintained and the

layout was adjusted to maximize usable space within the chassis.

1.4.3. Changes Made to Project Plan

No major changes have been made to our project plan. We have added a great deal of detail with

regard to funding and travel plans. We are on track with our timeline and have multiple potential

dates for full scale launches.


2. Vehicle Criteria

2.1. Design and Verification of Launch Vehicle

2.1.1. Flight Reliability and Confidence

The mission is to construct a Launch Vehicle capable of sending the Payload to an

apogee of 5,374 feet. The Launch Vehicle will then split into two subsections and use a

dual-deploy parachute system to land safely back on the ground. The Booster Stage will

split into the Lower Avionics Bay Section and the Booster Section when the main

parachute is deployed at 800 feet. The Payload Section will split into the Nose Cone

Section and Payload Section at 800 feet as well. For the mission to be successful there

must be minimal damage to the vehicle when the four subsections land. The vehicle must

also:

Reach an apogee of 5,280 feet

Be able to sit armed on the launch pad for 1 hour while remaining functional

Have a total impulse of no more than 5,120 Newton-seconds

Exit the launch rail with:

o A minimum static stability margin of 2.0 calibers of stability

o A minimum velocity of 52 feet per seconds

Deploy drogue parachute at apogee

Separate into the Booster Stage and the Payload Section at apogee and remain tethered

together

Successfully deploy a 48 inch diameter drogue parachute out of the Booster Stage after

initial separation, without the parachute tangling with the other section’s parachute

Booster Section separate into Payload Section, Lower Avionics Bay Section and Booster

Section that are tethered together also at 800 feet

Nose Cone Section separate from Payload Section at 800 feet and deploy 72 inch

parachute

Deploy a 72 inch elliptical main parachute that is connected to the two booster

subsections that are tethered together, slowing the velocity of the sections from 57.81 feet

per second to a final speed of 19.62 feet per second

At landing Launch Vehicle must:

o Sustain minimal damage to payload when landing

o Land in a 2500 foot radius from launch pad

o Land such that the launch vehicle can immediately launch again with minimal to

no repairs

o Safely house the electronics bay for data recovery

o Land at a kinetic energy no greater than 75 ft-lb

2.1.2 Design Alternatives from PDR

Nose Cone:

The design alternatives considered for the nose cone were all material alternatives. The

two main options that were being decided between were fiberglass and carbon fiber.


Fiberglass was chosen as the leading design because it is a tougher material than carbon

fiber. This means that it can endure more than carbon fiber before breaking. Fiberglass

also is not conductive and is radiolucent, meaning the material allows radiation to pass

through its walls which is especially helpful for allowing signals to pass to and from the

Launch Vehicle. This allows the altimeters to send the altitude data back to base.

Additionally, fiberglass is nearly half the price of carbon fiber and the carbon fiber would

not have been able to be machined by the team. Taking all of this into account the Nose

Cone will be made of fiberglass.

Payload:

The Payload section of the Launch Vehicle needs to house the rover payload. As with any

section of the Launch Vehicle, the material that comprises the body tube is crucial, and

difficult to decide upon. The first main alternative selection was the decision to use

carbon fiber instead of Blue Tube. The second decision was how to attach the parachute

to the Launch Vehicle.

As with other sections of the Launch Vehicle, fiberglass was the material ultimately

chosen. The other options that were considered were Blue Tube and carbon fiber. Blue

Tube was ruled out because it is not as durable as fiberglass and does not handle rough

weather conditions well. Carbon fiber was determined to be too expensive and difficult to

work with, so it was also decided against. In the end, fiberglass is the best choice for the

body tube of the Payload section because it is durable, inexpensive, and easy to work

with.

The attachment method for the parachute to the Launch Vehicle needs to be able to

ensure that the cords do not tangle during deployment. If the parachute cords were to

tangle during the separation, it would cause the parachute not to deploy and potentially

create a safety hazard while also harming the Launch Vehicle. The options for the

attachment method are an eye bolt, a U-bolt or a swivel hoist ring. The swivel hoist ring

was decided as the best option because it will ensure that the cords will be able to move

so would not tangle. The parachute will be connected to the aft bulkhead via a ¼-20

swivel hoist ring.

Booster Section:

The Booster Section of the Launch Vehicle will hold the motor. This means that the

section needs to be sturdy. The Booster Section is also where the fins are attached,

creating another challenge for design. The Motor Section will also use fiberglass with

carbon fiber and blue tube as the possible alternatives explored. The four fins will be a

trapezoidal shape rather than a triangular or elliptical shape.

As with other sections of the Launch Vehicle, fiberglass was ultimately chosen as the

material while Blue Tube and carbon fiber were considered. Blue Tube was ruled out

because it is not as durable as fiberglass and it does not last in rough weather conditions.

Carbon fiber was determined to be too expensive and difficult to work with, so it was

also decided against. In the end, fiberglass is the best choice for the body tube of the

Booster section because it is durable, inexpensive, and easy to work with.


Another alternative in the Booster section was the number of fins used. In the

OpenRocket simulation software, when the Launch Vehicle was tested with 3 fins, the

stability of the rocket was greatly reduced to the point that the design would not work.

Because of the stability issue, the number of fins on the Launch Vehicle was increased to

four fins. These fins increased the stability of the Launch Vehicle and allowed it to

function properly.

In addition to number of fins, both triangular and elliptical shapes were considered for the

fins, however both were decided against for different reasons. The triangular design was

decided against due to the fact that the Launch Vehicle will be colliding with the ground,

and the triangular design is not structurally sound enough and has a tendency to break.

Elliptical shaped fins were also decided against due to their increased difficulty in

machining. Trapezoidal shaped fins avoid both of these problems because they will be

less likely to break during landing, and the straight lines allow easy fabrication.

2.1.3 Computer Aided Design (CAD) Drawings

Figure 2.1.3.1 - OpenRocket Design of the Launch Vehicle

Figure 2.1.3.2 - CAD Dimensional Drawing of Launch Vehicle


Figure 2.1.3.3 - CAD Dimensional Drawing of Nose Cone Section

Figure 2.1.3.4 - CAD Dimensional Drawing of Payload Section


Figure 2.1.3.5 - CAD Dimensional Drawing of Avionics Bay Section

Figure 2.1.3.6 - CAD Dimensional Drawing of Booster Section

2.1.4. Discuss the Integrity of Design

When deciding the shape of the fin, it is crucial to consider the drag properties of the

various shapes. An ideal design for a fin would minimize drag while ensuring stability

and the lift of the Launch Vehicle. The fins also must be sturdy enough to withstand the


landing of the motor section of the Launch Vehicle. If the fins are not in a durable enough

they will break when landing. During the design of the fins, trapezoidal, elliptical, and

triangular fins were considered. The trapezoidal fins used provided lower drag than

elliptical fins, allowing the Launch Vehicle to reach its highest potential. In addition, the

tips of the four trapezoidal fins are effective at creating lift and provide stability in cases

where the path of the Launch Vehicle is disturbed.

Throughout the initial design process, great care was taken in deciding the materials to

make up the structural elements of the Launch Vehicle. A structure element can be

considered a load bearing element at least point throughout out the flight mission. The

main structural elements that were taken into consideration were the airframe, coupler,

bulkheads, fins, and the electronics bays. The airframe and coupler will be made of

composite G12 fiberglass from Mad Cow Rocketry. This material was chosen due to its

strength and stiffness as compared to alternatives such as Blue Tube. The G12 fiberglass

also has good impact resistance for landing on harder surfaces, for increased flight

mission success and re-launchability. The nose cone will also be made from G12

fiberglass from Mad Cow Rocketry for the same structural reasons. A similar material for

the fins was chosen as the fins of the rocket can often break on landings. G10 Fiberglass

was chosen for the fin material as it is a lightweight and thin fin material. While plywood

fins for the Northeastern University AIAA Student Chapter have been popular and

successful in the past, fins made of composite wood are susceptible to water damage, and

are often thicker with a rougher surface finish. G12 has similar mechanical properties to

G10 fiberglass, but also features more electrical resistance.

Table 2.1.4.1. - G10 Fiberglass Material Properties

Density 0.0650 lb/in3

Tensile Strength at Break 38000 psi Crosswise

45000 psi Lengthwise

Flexural Strength 65000 psi Crosswise

75000 psi Lengthwise

One of the most important components to design for is the bulkheads and centering rings.

The bulkheads receive a variety of tensile and compressive forces and large impulses

from ejection charges and parachute deployments. Centering rings transfer the forces

from the motor to the airframe on launch and also experience large force loads. The

selected material for is laser cut ¼ inch thick Mil-P-6070 aircraft mahogany plywood,

whose specifications are found in Table 2.1.4.2.


Table 2.1.4.2. MIL-P-6070 Aircraft Mahogany Plywood Material Properties

Material Type Laminate (Orthotropic)

Young's Modulus (Longitudinal) 1.39 - 1.7 x 106 psi

Young’s Modulus (Transverse) 0.119 - 0.133 x 106 psi

Tensile Strength (Longitudinal) 8.96 - 10.9 ksi

Tensile Strength (Transverse) 0.479 - 0.595 ksi

Shear Strength (Longitudinal) 1.35 - 1.65 ksi

Shear Strength (Transverse) 4.06 - 4.95 ksi

Flexural Strength (Longitudinal) 9.63 - 11.8 ksi

Flexural Strength (Transverse) 0.479 - 595 ksi

While typically not load bearing, the avionics bay can experience large forces throughout

the flight and should be designed with this in mind. The avionics bays will be made out

of ABS plastic, which has a higher melting point than PLA. Although they will be 3D

printed, which could reduce structural integrity; it provides freedom in the design of the

electronics bay as a whole. In addition, 3D printed, ABS electronics bays have been

proven to work in the past for Northeastern’s AIAA chapter

The motor being used is a 75 millimeter 4 grain Cesaroni reloadable motor type. The

reusable casing will be loaded into the aft end of the rocket into a 75 millimeter Blue

Tube motor mount tube. The motor casing will be inserted in its entirety into the Launch

Vehicle and secured into the motor mount tube with a motor retainer made of 6061-T6

aluminum with a corrosion resistant black anodize coating. The selected motor retainer is

produced by Aero Pack and has been shown to be successful in previous Northeastern

University AIAA Student Chapter launches. This is a very popular method of motor

retention due to its ease of use, and ability to quickly secure or remove a loaded motor

casing. The Aero Pack motor retainer will be secured to the Launch Vehicle using the

adhesive JB Weld. The motor retainer secures the motor in the axial direction and

distributes the thrust load on the rocket through the motor mount tube and centering

rings. The motor mount tube secures the motor radially and aligns the motor axially to the

airframe. The friction fit between the motor casing and the motor mount tube and the

motor retainer fixes the casing from rotating during the flight mission.


2.1.3. Justification for material selection, dimensioning, component placement, and

other unique design aspects

The outer body of the launch vehicle is made of fiberglass. The airframe, Nose Cone, and

coupler tube are G12 Fiberglass, and the fins are G10 fiberglass. This material was

chosen for its strength and durability. Although this material is also heavy and can cause

hazards when machined, it is less expensive than similarly heavy and hazardous materials

like carbon fiber. Fiberglass also is not conductive and is radiolucent--the material allows

radiation to pass through its walls--allowing signals to pass to and from the Launch

Vehicle. This allows the altimeters to send the altitude data back to base.

The Nose Cone of the Launch Vehicle is an ogive shaped fiberglass cone made by

Madcow Rocketry that is 30 inches in length. The Nose Cone is 0.079 inches thick, and is

attached to the top of the upper section of the body tube of the launch vehicle. The base

diameter of the Nose Cone is 6.17 inches, with a shoulder that gives the portion of the

Nose Cone that fits into the body tube (the bottom being 5.5 inches) a 6 inch diameter,

making it fit snugly into the body tube. The estimated weight of the Nose Cone section is

5.46 pounds, which is fairly heavy but also allows the Nose Cone to be tough and flex

more, making it less prone to breaking.

The Payload section is connected to the Nose Cone Section and the Avionics Bay section

by a 12 inch coupler. There will be 6 inches of the coupler in each of the sections with a

thickness of 0.104 inches. There will be a laser cut plywood bulkhead with a diameter of

5.792 inches and a thickness of 0.25 inches at the end of the section closest to the Nose

Cone. Next to the bulkhead will be the Payload, which will be 14 inches long and have a

diameter of 5.5 inches. The bulkhead will be 68 grams and the approximate weight of the

Payload is 2500 grams.

There will be another plywood bulkhead with a diameter of 6 inches and a thickness of

0.25 inches with a weight of 73 grams. On the other side of this bulkhead is the

Electronics Bay for payload deployment. The Electronics Bay weighs approximately

1500 grams will be 3D printed out of ABS plastic. The Electronics Bay is 8 inches long

with a diameter of 5.5 inches. The Electronics Bay is housed between two bulkheads, the

first of which was previously described. The second bulkhead will have a diameter of

5.792 inches, a thickness of 0.25 inches and a weight of 68 grams. The overall complete

length of the payload section of the Launch Vehicle will be 19 inches. There will be three

bulkheads total, with the Payload and the Electronics Bay in between them. The total

weight of the Payload Section will be 15.67 lb.

The parachute will be connected to the aft bulkhead via a ¼-20 swivel hoist ring. The

swivel hoist ring was chosen to decrease the probability that the parachute cords would

tangle. Design alternatives like the eyebolt or U-bolt could have led to the cords begin

tangled and caused the parachute to not deploy properly.

The avionics bay is a case located inside the fiberglass Booster Section of the rocket that

contains two StratoLogger altimeters, as well as a XBEE Pro XSC (S3) GPS. The mass


of the avionics bay is 750 grams. This bay is between two bulkheads of diameter 6 inches

and thickness 0.25 inches. The two bulkheads are 11 inches apart from each other. There

are two blast caps on each bulkhead, on the side facing out and away from the Electronics

Bay. There are also two terminal blocks on each side of the bulkhead. There are ⅜-16

inch forged eye bolts on each side of the bulkheads. The total mass of the lower avionics

bay section is 8.97 lbs.

The Booster Section is 36 inches long and has an inner diameter of 6 inches and outer

diameter of 6.17 inches. The material of the Booster Section is made of fiberglass. The

Booster Section also contains an inner tube with an outer diameter 3.1 inches. This inner

tube is made from Blue Tube and is attached to the body tube with four 0.25 inch birch

centering rings placed at 0, 8.25, 14.25, and 21 inches from the bottom of the rocket. The

section ends with a 12 coupler tube connecting the two lowest body tube sections. At the

end of the coupler tube, a 0.25 inch birch bulkhead separates marks the end of the booster

section. On this bulkhead, a half-inch 20 swivel hoist ring is in place to anchor the main

parachute to the booster section of the rocket. The estimated mass of the Booster Section

is 7.76 lbs with a loaded motor.

The Booster Section has 4 fiberglass trapezoidal fins. OpenRocket simulations

demonstrated that 4 fins would optimize the stability of the rocket. The trapezoidal design

is simple to machine and is a structurally sound shape for ground impact.

The motor mount of the launch vehicle is 24.25 inches long, has an inner diameter of

54mm inches, and an outer diameter of 3.1 inches. It is composed of Blue Tube. The

subsystem will house the Cesaroni L1115 Classic reloadable motor. Simulations

performed on the open source flight simulation software OpenRocket predict that this

motor will the launch the Launch Vehicle at an altitude of approximately 5302 feet.


2.2. Subscale Flight Results

2.2.1. Subscale Flight Results

Figure 2.2.1.1


Figure 2.2.1.2

Figure 2.2.1.3


Figure 2.2.1.4

The data gathering devices used during test launch were four StratoLogger CF altimeters. The

altimeters were located in the avionics bay and recorded a starting altitude of -1 meters, or -3.28

feet, and an apogee of 705 meters, or 2,313 feet. Thus, the recorded apogee, if the Launch

Vehicle had launched from 0 meters or feet, was 706 meters, or 2,316.28 feet. This is in

comparison to the Open Rocket flight simulation which had a predicted apogee of 2,723 feet. The

graphs of the recorded altitude over time are located above.

Scaling Factors The size of the subscale compared to the full size Launch Vehicle generally followed a ratio of

2:3. The figure below shows the design of the final full scale launch vehicle.

Figure 2.2.1.5


The inner diameter of the final Launch Vehicle was 6 inches and the inner diameter of the

subscale measured 4 inches. The length of the full scale Launch Vehicle will be 148 inches while

the subscale Launch Vehicle was a length of 108 inches. This deviated from the general rule to

ensure greater stability. The subscale model is located below.

Figure 2.2.1.6

The size changes were made to increase the safety factor of the design when testing the

experimental ejection system for the parachutes. The subscale Launch Vehicle had two

parachutes; a main parachute measured 48 inches and a drogue parachute measured 36 inches.

The full scale Launch Vehicle will include two parachutes; a main parachute measuring 72

inches, following the 2:3 ratio, and a drogue parachute measuring 48 inches. In both models the

drogue parachutes have 0.8 drag coefficients and the main parachutes have 2.2 drag coefficients.

The result of this is that the subscale design can more accurately represent the behavior of the

final design. The stability of the full scale Launch Vehicle was 3.35 caliber, and the stability of

the subscale Launch Vehicle was 1.76 caliber. The stability between both models was intended to

stay relatively constant so that the subscale Launch Vehicle and full scale Launch Vehicle would

behave in similar manners in similar launch conditions.

Launch Day Conditions The launch took place in Church Hill, Maryland on December 16th, 2017. The temperature then

ranged from a low of -5 degrees Celsius to a high of 7 degrees Celsius. There was a cover of

scattered clouds, but no precipitation. The wind speed was recorded at 9 miles per hour on the

ground, so the Launch Vehicle was launched at a 5 degree angle. A simulation using these

conditions can be found below.


Figure 2.2.1.7

2.2.2. Perform an Analysis of the Subscale Flight

Predicted Flight Model vs. Actual Flight Data

In preparation for the launch of the subscale model of the Launch Vehicle, taking place on

December 16th, the team generated simulation data to predict the successfulness of the launch

utilizing the Open Rocket software. That simulation can be seen below.

Figure 2.2.2.1: Open Rocket Simulation


In comparison, displayed below is the sub-scale’s flight data in the actual launch as recorded by

one of the Launch Vehicle’s stratologers during flight.

Figure 2.2.2.2: 12/16/17 Flight Data

As displayed in the figures above, the direct apogee of the Launch Vehicle may have been

slightly lower than estimated due to a number of possible factors such as slight differences in

weight, weather patterns, etc. The general path of descent, and velocity throughout its movement

were generally accurate to the simulated data, this can be observed by the similar slopes of the

two graphs. The test launch for the subscale in Church Hill, Maryland had an apogee of 705 meters, or 2,313

feet. The predicted apogee based on the Open Rocket software was 830 meters, or 2723 feet.

Potential causes for the difference in apogee may include weather conditions and launching at a 5

degree angle.

Estimate the drag coefficient of full scale rocket with subscale data. For the full scale rocket, the drag coefficient was estimated utilizing flow similarity. Due to the

parallels in the geometry of both rockets, the streamlines that go along the bodies of the launch

vehicles are also proportionally similar. Using this knowledge combined with the Buckingham Pi

Theorem, which relates fluid density flowing along the launch vehicle (ρ), dynamic viscosity (μ),

characteristic linear dimension (L), and fluid velocity (v), the team is able to define the

corresponding Reynolds number derived in the equation below,

This Reynolds Number is the ratio between inertia force and viscous force, and is able to be

utilized to characterize the flow's behavior, and thereby determine whether it is laminar,


turbulent, or in transition. Using this, the drag coefficient can be determined by relating the

velocity of the fluid flowing past the launch vehicle in the following equation,

Using this knowledge, the Open Rocket software can be used to measure and plot the Reynolds

numbers against the time of flight of the rockets, for both the subscale and the full scale rockets.

Figure 2.2.2.3: Subscale Reynolds vs. Time


Figure 2.2.2.4: Full Scale Reynolds vs. Time

While the maximum Reynolds numbers differ by nearly 100%, they share the same general flight

profile, shape, and slope, graphically, which shows that they share a common ratio and are

thereby proportional in nature. The proportionality between Reynolds Numbers was used to estimate the full scale’s Coefficient

with the subscale’s Coefficient as a reference point. Utilizing OpenRocket once more, coefficient

of drag can be estimated alongside altitude vs. time.


Figure 2.2.2.5: Subscale Coefficient of Drag

Focusing on the data pertaining to the drag coefficient leading up to recovery device deployment,

the coefficient of drag is to be estimated an average of roughly 0.604125. With this the coefficient

of drag of the full scale can be estimated to be relatively close.

2.2.3. Impact of Subscale Flight Data on Full Scale Launch Vehicle

Design choices made in the subscale version of the Launch Vehicle have influenced future

decisions that apply to the full scale Launch Vehicle. The general design of the launch vehicle

and recovery systems were reliable and should be implemented into the full scale. However, the

subscale flight data has had a slight impact on the full scale. In the first launch of our subscale

launch vehicle, the ejection charge responsible for splitting the rocket at the junction between the

payload and booster sections deployed, but failed to separate the sections. In the second subscale

launch, this issue was fixed by increasing the charge and using 2 shear pins instead of 4 shear

pins in the coupler between those two sections. This second launch proved this change to be

effective, so it will be utilized in the full scale design. Another change to the full scale design will

be organization of the electronics bays. Wire management needs to be improved for organization

and spatial optimization. In addition, the plastic structure needs to be revised as a result of issues

with battery placement and due to increased space in the full scale. Lastly, due to issues with the

labeling of the different body tube sections, there was confusion with how each section was

oriented with one another. In the full scale, labeling will be more precise and easier to understand

in order to avoid this confusion. Overall we have been able to learn from the subscale. We have

implemented changes and as a result the full scale Launch Vehicle will be better off.


2.3. Recovery Subsystem

Chose recovery design alternative from PDR

Figure 2.3.1: Full Scale Flight Profile

After reviewing the three main design alternatives for the recovery system, the original design was

chosen, with a few changes. This design has the Launch Vehicle falling in two pieces: the payload

section and the booster. At apogee, the booster section separates with a black powder charge, deploying a

48 inch drogue. At 800 feet, a black powder charge separates the payload section and nose cone section,

deploying the 72 inch, lower main parachute. Shortly after, the payload section and lower avionics bay

section separate and begin falling no longer tethered. In addition, a 72 inch main parachute for the lower

two sections is released during this separation. The separation between the nose cone and payload

sections occurs first such that there is no section in free fall during the main parachute deployment

process. Despite minor changes in timing of separations and exact separation points, this design mirrors

the original recovery system design. It was chosen over the other two alternatives for one major reason;

the risk of the parachutes becoming tangled before or as they are pulled out of the Launch Vehicle. The

first alternative involved using a tender descender in order to house both booster sections in the same

compartment. Although this would reduce the number of falling sections, it increases the chance of the

parachutes becoming tangled when the Launch Vehicle separates, simply because they are in the same

compartment. The second alternative was similar to the chosen design with one notable exception. Due

to the placement of the black powder charges, the parachutes would be pushed further into the Launch


Vehicle before being jerked out by the separation. Even though this design requires fewer sections of

coupler tube, it greatly increases the chance the parachutes become caught, preventing them from

deploying correctly. Therefore, the third and original design alternative was chosen, as it presented the

least risk of parachute entanglement which would result in failure of the recovery system as a whole. Parachute, harnesses, bulkheads, and attachment hardware

Parachutes

Figure 2.3.2: Iris Ultra 72” Central Parachute

Manufacturer Fruity Chutes

Diameter 72 Inches

Type Standard Nylon Toroidal (Annular)

Drag Coefficient 2.2

Rating 28lb @20fps

Material Nylon

Both main parachutes are Iris Ultra Light Parachutes manufactured by Fruity Chutes. The parachutes for

both main sections are the same parachute, meaning they have the same diameter, drag coefficient, and


rating. By this metric, both parachutes have a 72 inch diameter, a rating of descent of 20 feet per second

at 28 pounds of weight.

Figure 2.3.3: Drogue Parachute

Manufacturer Sunward Group Ltd

Diameter 48”

Type Octagonal

Drag Coefficient 0.8

Rating 65lb @22fps

Material Nylon

The drogue parachute located in the lower booster main has a 48 inch diameter and unlike the main

parachutes which have toroidal shapes, has an octagonal shape instead. With a lower drag coefficient of

0.8 compared to the main parachutes of 2.2, it is able to maintain a higher rating of 22 feet per second for

65 pounds. Bulkheads The bulkheads featured on the full-scale design of the Launch Vehicle are crafted from ¼ inch birch

plywood, and total to eight throughout the length of the Length Vehicle. All bulkheads have a diameter of

5.787402 inches, save for the bulkhead nearest the nose cone, and the bulkhead nearest the motor section,

which each have a diameter of 5.984252 inches.


Eyebolt

Figure 2.3.4: Eye bolt

All three parachutes are attached utilizing ⅜-16in eyebolts. These eye bolts each have a vertical weight

capacity of 1300 pounds and as such are sufficient for successfully linking the parachutes to the rest of the

Launch Vehicle. Shock-Cord

Figure 2.3.5: Shock-Cord Lengths

Within the Launch Vehicle, both main parachutes are connected via a ½ inch thick Kevlar shock-cord,

and the drogue parachute is connected via a ¼ inch Kevlar shock-cord. Each of the three shock-cords

differs in length, with the upper main’s cord being 33 feet in total, divided as 21 and 12 feet, the lower


main being 18 feet in total, and the drogue being 22 feet in total, divided as 9 and 13 feet. Each shock

cord is connected via eyebolts, as detailed above.

Electrical Components

The Launch Vehicle will have in total six Perfectflite StratoLogger CF altimeters. Each individual

pressure-based altimeter is capable of powering two deployments, containing internal power-switches

allowing for an external system to turn the altimeter on and off at will. Thus far, the team has found

through experience that the StratoLogger CF altimeter is functional and extremely reliable, given that the

majority of Northeastern AIAA Launch Vehicles have utilized either the StratoLogger or StratoLogger

CF altimeter.

Figure 2.3.6: StratoLogger (Perfectflite.com)

Table 2.3.7: StratoLogger CF Specifications

Power 4V - 16V, nominal 9V battery

Current Consumption 1.5 ma

Launch Detect 160’ to 300’ AGL, default 160’

Main Deploy Altitude 100’ AGL to 9,999’ AGL

Maximum Altitude 100,000’ MSL

Altitude Resolution 1’ up to 38,000’ MSL < 2’ to 52,000’ MSL < 5’ to 72,000’ MSL

Measurement Precision +/- (0.1% reading + 1 foot) typical

Flight Data Logged Altitude, temperature, battery voltage

Recording Time Per Flight Over 18mins

Operational Temperature -40C to +85C (-40F to +185F)

Sample Rate 20 samples per second

Cost $54.95


In order to provide the safest recovery of the Launch Vehicle, there needs to be redundancy in the

recovery system. For every ejection event, an e-match will spark and ignite a small black powder charge

to separate the rocket. There will be two Stratologger altimeters responsible for every separation event.

Two Strataloggers will be located in the Lower Avionics Bay for the separation at apogee. There will be

another two Stratologgers in the same Avionics Bay and the Nose Cone Avionics Bay for the two

separation events that occur at 800 feet on the descent. Traditionally, one Stratologger is able to provide

enough power for two ejection events. By having six Stratologgers perform the job that can be performed

by two; we are ensuring that each altimeter will have sufficient current to ignite the e-match and record

data for the entirety of the flight. In terms of power-sources, each Altimeter utilizes a nine volt battery, specifically, the Duracell Quantum

Series 9-Volt Battery. After having compared this battery to several alternatives, including the Energizer

Lithium Battery, the Amazon Brand 9-Volt Batteries, the team determined the Duracell model to be the

best battery. In making this choice, the team considered capacity, battery construction, and internal

resistance. The Duracell was chosen following these specifications, and those mentioned in the below

table:

Figure 2.3.8. : Duracell Quantum 9V

Table 2.3.9. Duracell Quantum 9V Specifications

Manufacturer Duracell

Nominal Voltage 9.6 V

Capacity 4700

Impedance 1.7 ohm

Mass 46 grams


Figure 2.3.10: Sample Recovery System Circuit Diagram

Figure 2.3.11: XBee PRO S3B

The tracking system used is contained within the payload sections and lower avionics sections of the

launch vehicle and operates at a frequency dependent on the electronic systems. The factors that this

frequency is dependent on include the polling rate of the GPS, the rate at which the Arduino runs the

code, and the rate at which the XBee transmits data. The limiting factor of these processes is the 10Hz

rate at which the GPS polls for new data. Therefore, we will be receiving location updates at a rate of

approximately 10 times per second.


2.4. Mission Performance Predictions Flight Profile Simulations

Figure 2.4.1: Flight profile simulation.

20 simulation runs on OpenRocket found an expected apogee of 5,303 feet, plus or minus 6 feet.

MATLAB simulations find slightly lower values: 17 s for time to apogee and 4700 ft for apogee.

Table 2.4.2: Component Mass Distribution

Component Mass

Nose Cone Section

Body Tube 1182 g

Ogive Fiberglass Nose Cone 1043 g

Bulkhead 73 g

Nose Cone Avionics Bay 200 g

Shock Cord 3.29 g

Parachute 50.6 g

Payload/Booster Section

Body Tube 936 g

Body Tube 2217 g


Body tube 1478 g

Bulkhead x3 68 g each

Bulkhead x4 73 g each

Centering ring x4 53.5 g each

Tube coupler x3 697 g each

Shock Cord x2 4.11 g each

Payload Mechanical 2250 g

Payload Deployment Electronics Bay 1000 g

Booster Section Avionics Bay 1000 g

Parachute Booster Main 181 g

Parachute Booster Drogue 81.5 g

Trapezoidal fin set 1066 g

Inner Tube 248 g

Figure 2.4.3: L1115 Thrust Curve.


Stability Margin Stability: 3.35 caliber CP: 7.74 feet from tip of nose cone CG: 9.45 feet from tip of nose cone

Figure 2.4.4: Launch Vehicle Model.

Kinetic Energy calculations Kinetic energy was calculated by determining the terminal velocity of each individual section of the

Launch Vehicle and the point in its descent when its speed is not changing due to acceleration. At

terminal velocity, the net force any individual section of the launch vehicle experiences amounts to zero.

The main forces involved in this net force calculation are the force of gravity and the force of the wind

drag on any given section. These force functions are given by the following equations.

FD=12CDAVT2

FG=mg

Where FD is the Force of Drag, ⲣ is the constant density of air (1.225 kilometers per cubic meter), CD is the

Coefficient of Drag which varies for each parachute, A is the surface area of the exposed surface, VT is the

terminal velocity, m is the mass of the given section, and g is the gravitational constant for earth (32.174

feet per seconds squared). Utilizing the relationship of Newton’s Second Law, which sets the sum of all

forces acting on an object equal to its mass times its given acceleration. As the object reaches its terminal

velocity and stops accelerating, the sum of forces is equal to 0. Using this, the team was able to calculate

the terminal velocity for every section.

FD-FG=ma FD-FG=0

12CDAVT2-mg=0

VT=(mg12ⲣCDA)1/2

Knowing the terminal velocity allows for calculation of Kinetic Energy as follows, where V is substituted

for the equation above, as follows:

Ek=mVT22

Ek=m2g24ⲣCDA


Using the Kinetic Energy equation listed above, the team began calculations based on the relevant

parachute data and section. Because the relevant data for the force of drag in the kinetic energy equation

above was related to the parachutes, the table is organized by parachute instead of section.

Table 2.4.5: Parachute Coefficients and Area

Parachute CD A(in2)

Nose Cone Main 2.2 6840

Drogue 0.8 1564.4

Booster Main 2.2 6840

Given the parachute data and the mass of each section, the team was able to calculate the Kinetic Energy

of the rocket. The data for these calculations are listed in the following table, which is organized by

section, mass, parachute, and Kinetic Energy.

Table 2.4.6: Kinetic Energy Data by Section

Section Mass Total (lb) Parachute Data (CD * A) Kinetic Energy (ft * lb)

Nose-Cone 2.86 Nose Cone Main 21.8

Payload 15.67 Drogue + Booster Main 62.4

Lower Avionics Bay 6.50 Drogue + Booster Main 33.4

Booster 7.75 Drogue + Booster Main 45.8

The team was conscious of the maximum Kinetic Energy for any given section, 75 foot-pounds, and

attempted to obey that constraint. As observed in the table above, all sections stay within this limit.

Wind Drift Calculations Drift due to wind can conservatively be calculated by simulating a flight profile to find flight time and

multiplying by the wind speed. Under a nominal flight configuration, the rocket would experience the

following drifts:


Table 2.4.7: Launch Vehicle Drift Calculations Using OpenRocket Simulation Software, 10 Simulation

Average

Wind Speed (mph) Drift (ft)

0 9

5 525

10 1,125

15 1,782

20 2,490

Table 2.4.8: Launch Vehicle Drift Calculations Using Hand Calculations

Wind Speed (mph) Flight Time (s) Drift (ft)

0 142 0

5 141 1,034

10 156 2,288

15 158 3,476

20 156 4,576

The drifts show above are of the tethered lower avionics bay and booster sections, as these two lower

sections have a slower descent velocity as compared to the upper two sections, and will experience a

further drift. To calculate the total lateral drift of the rocket, simulations were performed in OpenRocket

and by hand. The OpenRocket simulation results are the average of 10 simulations. Using the total

descent time of the rocket from the terminal velocity of the Launch Vehicle under drogue and main

parachute and using the various wind speeds, the team calculated the total lateral drift using the following

equation:

Lateral Drift = Vw*t Where “t” is the total descent time and “Vw” is the wind speed. Calculations of the total lateral drift were

performed for the following wind speeds: 0, 5, 10, 15, and 20 miles per hour. These calculated values for

lateral drift are shown in the table above. For these hand calculations, it is assumed that the Launch

Vehicle does not drift on its ascent and its descent begins directly above the launch location. In reality this

is not the case as the Launch Vehicle will drift into the wind on its ascent due to the fins. The simulations

can take this factor into account while the hand calculations do not. For this reason, the hand calculations

can be considered to be conservative. For wind speeds of 15 and 20 mph, the Launch Vehicle under a nominal flight configuration would

exceed the maximum 2500 ft recovery in the hand calculations. For these speeds, the Launch Vehicle

would need a different descent profile that would allow the launch vehicle to spend less time in the wind,


and thereby have a lower lateral drift. In these high wind environments, it would be safer to deploy the

two main parachutes later in the descent profile in order to decrease drift. In such a scenario, the

stratologgers would be reprogrammed to deploy parachutes at a minimum of 400 feet above ground level

instead of 800 feet. The results of this change can be seen in the table below.

Table 2.4.9: Wind drift under high-wind flight configuration

Wind Speed (mph) Flight Time (s) Drift (ft)

15 126 2,772

20 123 3,607


3. Safety

3.1. Launch Concerns and Operation Procedures Final assembly and launch checklist

Note: All energetics including black powder and the motor will be handled by team mentors (Andrew

Goldstone, Faculty Advisor and Robert DeHate, NAR Mentor) as dictated by NASA USLI handbook

rules

Table 3.1.1: Final Assembly and Launch Safety Checklist

Step # Step Done Safety Officer

Verified

Night Before Launch

1 Measure black powder, mark, and store safely

2 Tie all knots (bowline)

3 Put batteries in electronics bays

4 Verify continuity through all recovery electronics

5 Check the stratologger presets

Launch Day

1 Connect wired between electronics bays and bulkheads

2 Insert electronics bays into couplers

3 Secure electronics bays with nuts on threaded rods on outsides of

bulkheads

4 Load black powder blast caps

4.1 Connect e-matches between terminal blocks with bulk in blast

cap

4.2 Put pre-measured black powder into blast caps

4.3 Tightly pack ejection wadding into blast caps so that it is tightly

packed and the blast cap is full

4.4 Tape over blast caps with masking tape to secure

5 Set screw electronics bays/coupler to body tube (size 8 screws)


6 Fold Parachutes

6.1 Iris Chute

6.1.1 Fold each color in half on itself with so that leads are off one side

and the opening is on the top with only one color visible

6.1.2 Fold in thirds

6.1.3 Fold in half

6.1.4 Z fold

6.2 Standard Parachute

6.2.1 Fold in half so that leads are on one side

6.2.2 Ensure leads are not tangles

6.2.3 Fold in thirds twice

6.2.4 Fold in half

6.2.5 Z fold

7 Wrap parachutes in thermal wadding

8 Pack parachutes in body tube

9 Shear pin together the separable sections (size 2 screws)

10 Put the motor in the casing and screw it closed

11 Put the casing in the motor mount

12 Secure the motor with motor retainer

13 Get cleared by the RSO

14 Go to the launchpad with rocket

15 Lower the launch rail with the knobs

16 Put rocket on to the launchpad using the rail buttons


17 Secure rail at 5°, away from controllers and spectators

18 Turn on recovery electronics with key switches

19 Listen to Stratologgers beeps to double check settings including

deployment altitude, voltage of battery, and continuity

20 Insert ignitor all the way up motor

21 Secure igniter wires with the provided motor cap

22 Connect alligator clips such that they do not touch

23 Check continuity

24 Move designated distance away for launch

25 Launch

26 Track all falling sections during recovery

27 Collect rocket components after RSO has opened the range

If any step is missed a critical failure in the launch could occur. Each step must be checked by the team

safety officer in addition to the personnel who performed the task. Failure to comply with all recover

related steps could result in an object in free fall.

Recovery Preparation

Failure of recovery systems can cause a major safety risk because if something goes wrong the launch

vehicle could fall without parachutes. This is a major safety risk and therefore careful consideration has to

be paid when assembling the recovery systems. To mitigate this risk every step on the above list will be

followed with care and the safety officer and other members of the team will inspect the vehicle

throughout construction for safety risks.

The individual responsible for maintaining the safety, quality, and procedure checklists for the recovery

preparation is Rebecca Holleb.

Motor Preparation

The person packing the motor into the launch vehicle should be wearing gloves in order to protect their

hands.

1. Inspect the motor casing for defects

2. Inspect the motor for damage

3. Follow the instructions for assembly as stated by the manufacturer

Note: this list is only safety related steps for setup.


A faulty motor could cause a safety issue because it could partially burn and fall too soon without

deploying parachutes, it could explode on the launch pad, or it could explode in the air over bystanders.

This can be mitigated by inspecting the motor and the casing for defects before launching.

The individual responsible for maintaining the safety, quality, and procedure checklists for the motor

preparation is Rebecca Holleb.

Setup on Launcher

The risks to the individuals setting the launch vehicle up on the launch pad are centered on the parachute

charges. The armed e-bays could set off the black powder charges, however the chances are slim since the

e-matches will not go off until the altimeters register a specified altitude.

The individual responsible for maintaining the safety, quality, and procedure checklists for the setup on

the launcher is Rebecca Holleb.

Igniter Installation

The biggest risk to individuals undertaking the igniter installation is the accidental ignition of the motor.

We will be instructed when to set up our rocket on the pad by the RSO and it is his or her responsibility to

make sure that while we are installing the igniter that no signal to launch be received by the launch pad

while we are setting up. The risk of such a thing happening is extremely low and as we have never heard

of or witnessed such an event; we consider the risk acceptable. There are no environmental risks related to

igniter installation.

The individual responsible for maintaining the safety, quality, and procedure checklists for the igniter

installation is Rebecca Holleb.

Troubleshooting

Below is a list of systems that could require troubleshooting. Note this is an abbreviated version focused

on safety.

1. Xbee: this system is important for vehicle recovery and data collection, yet is not crucial for

safety

2. Sensor Data: again important for data collection yet not crucial to safety

3. Electronics Bay: failure to troubleshoot if armed incorrectly could result in failure of parachutes

to deploy

If there is an electronics failure it could pose a risk to bystanders if the parachutes do not deploy. This will

be mitigated by checking continuity in the wiring before assembly and listening for the altimeter beeps

when the keys are turned closing the circuit. If there is a problem with the electronics the day of the

launch the vehicle will not be launched until the problem is solved.

The individual responsible for maintaining the safety, quality, and procedure checklists for

troubleshooting is Rebecca Holleb.

Post Flight Inspection


1. Make sure there is no un-detonated black powder. Failure to complete this step successfully

could result in serious injury due to potential unplanned detonation.

2. Check for any shrapnel or hazardous parts of rocket. Failure to complete this step could result in

injury.

Possible safety issues in retrieving the launch vehicle include possible broken material that could result in

injuries such as cuts as well as undetonated black powder charges that could pose serious safety

risks. The mitigation for these safety risks are that the vehicle will be inspected for both hazards before

being moved. If hazards are found they will be treated accordingly.

The individual responsible for maintaining the safety, quality, and procedure checklists for post-flight

inspection is Rebecca Holleb.

3.2. Safety and Environment (Vehicle and Payload)

Personnel Hazard Analysis

Table 3.2.1 : Personnel Hazard Analysis

Section Hazard Effects Causes Sever

ity

Probabi

lity

Mitigatio

n

Verificati

on

Launch

Hazards

Prematu

re

Motor

Ignition

Severe

burns/bodil

y harm,

general fire

hazard

Improper

storage and

handling of

motor.

Motor placed

near heat

source.

4 1 Follow

MSDS

storage

requireme

nts. Keep

away

from

ignition

sources

Use a

written

procedure

in

accordanc

e with

MSDS

requireme

nts to

ensure

safe

storage.


Explosi

ve

Motor

Failure

on

Launch

pad

Severe

burns/bodil

y harm,

general fire

hazard

Defective or

damaged motor.

4 1 Maintain

personnel

a safe

distance

away

from

launchpa

d (300-

500 ft

according

to NAR

High

Power

Rocket

Safety

Code).

Follow

MSDS

storage

requireme

nts. Insert

and

prepare

motor

correctly.

Written

procedure

for launch

will

include

proper

preparatio

n and

insertion

of motor.

Motor

fails

to

ignite

If proper

safety

procedures

are followed

there

is no

safety risk

Improper

connection

between motor

and igniter,

defective

motor, or wet

motor.

0 1 Follow

NAR

safety

procedure

(60

second

wait)

before

accessing

Launch

Vehicle

following

motor

failure to

ignite.

Check

connectio

ns prior to

launch.

Insert and

Include

NAR

safety

procedure

s in

launch

checklist.


prepare

motor

properly

before

launch

sequence

is

initiated.

Motor

will be

inspected

before

launch.

Unstabl

e

Flight

Path

The Vehicle

goes on a

flight path

that was

unanticipate

d

One or more

fins fall off,

unintended

oscillation of

the rocket as a

result of

dislodged

internal

components, or

high winds

during launch

4 1 Maintain

personnel

a safe

distance

away

from

launchpa

d (300-

500 ft

according

to NAR

High

Power

Rocket

Safety

Code).

Verify

internal

compone

nts in

electronic

s bays

and

payload

are

secure.

Do not

launch in

Include

NAR

specificati

on and

compone

nt

verificatio

n in the

launch

checklist.


high wind

condition

s.

Total

Recov

ery

Systems

Failure

No

parachutes

are

deployed,

the Vehicle

is in freefall

The charges do

not detonate or

the sections do

not separate to

release the

parachutes.

4 1 Confirm

that

quantity

of black

powder in

ejection

charges is

sufficient

to

separate

launch

vehicle

sections,

check

recovery

systems

in

electronic

s bay for

proper

connectio

ns, and

verify

altimeter

performa

nce.

Maintain

personnel

a safe

distance

away

Include

extensive

prelaunch

review of

recovery

systems.


from

launchpad

(300-

500 ft

according

to NAR

High

Power

Rocket

Safety

Code).

Partial

Recov

ery

Systems

Failure

Either the

drogue

parachutes

or the main

parachute

is not

deployed,

the Vehicle

descends

faster than

anticipated

The charges do

not separate

some of the

sections, the

sections do not

release the

parachutes, or

the parachutes

get

tangled on

release. Drogue

parachute or

main parachute

fails to deploy.

3 2 Confirm

that

quantity

of black

powder in

ejection

charges is

sufficient

to

separate

launch

vehicle

sections,

check

recovery

systems

in

electronic

s bay for

proper

connectio

ns, and

verify

altimeter

performa

Include

extensive

prelaunch

review of

recovery

systems.


nce.

Confirm

the

parachute

is packed

correctly.

Maintain

personne

l a safe

distance

away

from

launchpad

(300-500

ft

according

to NAR

High

Power

Rocket

Safety

Code).

Shock

Cord

Failure

The Vehicle

would not

be

connected

to

the parachu

te(s) and

would either

descend

faster than

anticipated

or be in

freefall

Damage or

defect to the

shock cords

connecting

the drogue

and/or main

parachute to the

vehicle. Shock

cord improperly

secured to

bulkhead.

4 1 The final

assembly

checklist

will be

followed

and the

shock

cord will

be

inspected.

Include

shock

cord

inspection

in launch

preparatio

n

checklist.


Construct

ion

Hazards

Power

Tool

Injury

Injury

incurred

while using

a power tool

Improper

training or tool

maintenance.

Human

error

4 2 Properly

train team

members

on power

tool

handling

, wear

proper

Personal

Protective

Equipmen

t

according

to each

power

tool’s

operator’s

manuals.

Distribute

guidelines

on proper

tool

handling

and PPE.

Tool

Injury

Injury

incurred

while using

a power tool

Improper

training or tool

maintenance.

Human

error

4 2 Properly

train team

members

on tool

handling,

wear

proper

PPE, and

ensure

First Aid

equipmen

t is

available.

Distribute

guidelines

on proper

tool

handling

and PPE.


Chemic

al

Hazar

ds

Injury

incurred

while using

chemicals

Improper

training or

equipment

maintenance H

uman error

4 2 Train

team

members

on

chemical

handling,

and

follow

proper

storage

requireme

nts listed

on

MSDS,

provide

proper

PPE

accordin

g to the

material’s

instructio

ns, utilize

chemicals

only in

areas

designat

ed for

their use.

Distribute

chemical

handling

and

storage

guidelines

in

accordanc

e are

MSDS

requireme

nts.

Fire

Hazard

Injury

incurred due

to a fire

Improper

training. Huma

n error

4 2 Keep fire

hazardous

materials

stored

properly

accordin

g to

MSDS.

Use

written

procedure

in

accordanc

e with

MSDS.


Failure Modes and Effects Analysis (FMEA)

Table 3.2.2 : Failure Modes and Effects Analysis (FMEA)

Item Failure

Mode Effects Causes Severity Probability Mitigation Verification

Laun

ch

Vehic

le

Drogue

parachute

fails to

deploy

Launch Vehicle

will not decelerate

prior to main

chute

deployment. Incr

eased risk of main

parachute failure,

loss of payload,

and personnel

injury.

Improper

packing

of drogue

parachute

causes

tangling

upon

deployme

nt. Impro

per

deployme

nt charge

used

destroys

parachute

.

2 1 Main parachute

will deploy even

if drogue

parachute

fails. Main

parachute and

shock cord are

able to take loads

of main

parachute

deployment

without drogue.

Charges and

packing will be

checked prior to

launch.

Follow a

checklist to

ensure that

the drogue

parachute is

packed

correctly.

Have several

people

recheck to

pack the

parachutes

correctly.

Main

parachute

fails to

deploy

Launch Vehicle

will descend

under drogue

parachute, faster

than

nominal. Likely

loss of payload

and increased risk

to personnel.

Improper

packing

of main

parachute

causes

tangling

upon

deployme

nt. Impro

per

deployme

nt charge

used

destroys

parachute

.

3 1 RSO will alert

crowd to Launch

Vehicle with

parachute failure,

Launch Vehicle

will be aimed

away from

spectators, and

all spectators will

be made aware

of launches

occurring.

Charges and

packing will be

checked prior to

launch.

Monitor

communicati

ons with the

Launch

Vehicle to

know

exactly

when the

failure

occurs to

allow for

maximum

warning

time. The

parachute

will be

checked

several times

to ensure

correct

packing.

Both

main/dro

gue

parachute

s fail to

deploy

Launch Vehicle

will be in freefall.

Total loss of

launch vehicle

and substantial

risk to personnel.

Improper

packing

of main

and

drogue

parachute

causes

4 1 RSO will alert

crowd to launch

vehicle with

parachute failure,

Launch Vehicle

will be aimed

away from

Monitor

communicati

ons with the

Launch

Vehicle to

know

exactly


tangling

upon

deployme

nt. Impro

per

deployme

nt charge

used

destroys

parachute

s.

spectators, and

all spectators will

be made aware

of launches

occurring.

Charges and

packing will be

checked prior to

launch.

when the

failure

occurs to

allow for

maximum

warning

time. The

parachute

will be

checked

several times

to ensure

correct

packing.

Shock

cord

failure

(drogue)

Parachute

detaches from

Launch

Vehicle. Increase

d loads on main

parachute

deployment. Incr

eased risk of main

parachute failure,

loss of payload,

and personnel

injury.

Weakeni

ng or

damage

to shock

cord from

accidenta

l cutting

or epoxy.

2 1 Shock cord will

be kept away

from

unintentional

cutting and

epoxy. Main

parachute and

shock cord are

able to take loads

of main

parachute

deployment

without drogue.

Shock

cord

failure

(main)

Parachute

detaches from

Launch

Vehicle. Launch

Vehicle will

descend under

drogue parachute

at increased

speed. Likely

loss of payload

and increased risk

to personnel.

Weakeni

ng or

damage

to shock

cord from

accidenta

l cutting

or epoxy.

3 1 Attachment point

is

reinforced. Shoc

k cord is ½”

Kevlar. Finite

element analysis

of shock cord

and bulkhead

performed to

ensure they can

withstand

parachute

deployment. Sho

ck cord will be

kept away from

unintentional

cutting and

epoxy.

Shock

cord

failure

(both)

Both parachutes

detach from

Launch Vehicle.

Launch Vehicle

will be in freefall.

Total loss of

launch vehicle

Weakeni

ng or

damage

to shock

cord from

accidenta

l cutting

4 1 Shock cord will

be kept away

from

unintentional

cutting and

epoxy. Shock

cord will be


and substantial

risk to personnel. or epoxy. inspected prior to

launch for

damage or

defects. RSO

will alert crowd

to launch vehicle

with parachute

failure, Launch

Vehicle will be

aimed away from

spectators, and

all spectators will

be made aware

of launches

occurring.

Ejection

charges

fail to

ignite

Launch Vehicle

will not separate

for parachute

ejection and will

fall in a ballistic

trajectory. Total

loss of Launch

Vehicle and

substantial risk to

personnel.

Improper

connectio

n to

charges,

improper

installatio

n of

charges,

defective

charges,

and

altimeter

improperl

y

calibrated

.

4 1 Confirm black

powder in

ejection charges

is sufficient to

separate Launch

Vehicle sections,

check recovery

systems

electronics bay

for proper

connections, and

verify altimeter

calibration.

Prematur

e

detonatio

n of

ejection

(charges

(in

Launch

Vehicle)

Premature

separation of

vehicle

stages. Rapid

unplanned

disassembly of

vehicle.

Improper

calibratio

n of

altimeter.

Imprope

r

handling

of e-

matches.

4 1 Follow MSDS

storage

requirements. C

harges will be

inspected during

installation. Alti

meter

functionality will

be verified

before it is

connected to

charges.

Partial

deployme

nt of

drogue

parachute

Launch Vehicle

will decelerate

less until main

chute

deployment. Incr

eased risk of main

parachute failure,

loss of payload,

and personnel

Drogue

parachute

becomes

tangled

upon

deployme

nt. Impro

per

packing.

3 2 Main parachute

will deploy even

if drogue

parachute

fails. Main

parachute and

shock cord are

able to take loads

of main


injury. parachute

deployment

without drogue.

Packing will be

checked prior to

launch.

Partial

deployme

nt of

main

parachute

Launch Vehicle

will descend

under drogue

parachute and

partially deployed

main, faster than

nominal. Likely

loss of payload

and increased risk

to personnel.

Main

parachute

becomes

tangled

upon

deployme

nt. Impro

per

packing.

3 2 Launch Vehicle

will be aimed

away from

spectators, and

all spectators will

be made aware

of launches

occurring.

Packing will be

checked prior to

launch.

Partial

deployme

nt of both

drogue/m

ain

parachute

s

Vehicle will fall

at increased speed

under partially

deployed

parachutes.

Potential loss of

launch vehicle

and substantial

risk to personnel.

Main and

drogue

chutes

could

become

tangled

with one

another.

Improper

packing.

3 1 RSO will alert

crowd to launch

vehicle with

parachute failure,

Launch Vehicle

will be aimed

away from

spectators, and

all spectators will

be made aware

of launches

occurring. Packi

ng will be

checked prior to

launch.

Rapid

unplanne

d

disassem

bly of

Vehicle

(RUD)

Vehicle will break

apart in an

unplanned

manner. Loss of

vehicle structural

integrity and

increased risk to

personnel due to

vehicle

components in

freefall.

Structural

defect in

Launch

Vehicle

body.

4 1 In the event of a

RUD with the

vehicle breaking

apart along its

coupler tubes,

parachutes may

passively deploy

for each

section. Verify

that the Vehicle

is structurally

stable.

Explosive

motor

failure on

launch pad

Significant risk of

severe personnel harm

and fire

hazard. Significant risk

of loss of vehicle and

payload.

Defective or

damaged motor. 4 1 Follow MSDS storage

requirements. Insert and

prepare motor properly

before launch sequence is

initiated. Motor will be

inspected before launch.


Launch

Operations

Premature

motor

ignition (on

launch rail)

Significant risk of

personnel harm, fire

hazard, and

uncontrolled

launch. Potential loss

of vehicle and payload.

Improper storage or

handling of

motor. Improper

connection between

motor and igniter.

4 1 Maintain personnel a safe

distance away from

launchpad (300-500 ft

according to NAR High

Power Rocket Safety

Code). Follow MSDS

storage requirements. Insert

and prepare motor properly


initiated.

Motor fails

to ignite Vehicle will not launch. Improper connection

between motor and

igniter Motor igniter

falls out. Defective

motor. Motor is wet.

0 1 Follow NAR safety

procedure (60 second wait)

before accessing Launch

Vehicle following motor

failure to ignite. Check

connections prior to launch.

Insert and prepare motor

properly before launch

sequence is initiated. Motor

will be inspected before

launch.

Lithium Battery is

damaged to point

of failure (in

Rocket)

Possible fire or

explosion of

payload and

launch vehicle

Any form of

sharp object

jarred during

course of flight

4 2 Instructions for handling will be

followed as well as protection of

battery within electronics bay

implemented.


Payload

Lithium Battery is

damaged to point

of failure (on

ground)

Possible fire or

explosion of

payload

Any form of

sharp object

dislocated upon

landing

2 2 Instructions for handling will be

followed as well as protection of

battery within electronics bay

implemented.

Punctured

pressurized gas

containment

Leak and

possible

explosion

Any jagged

edges 3 2 All guidelines set forth by NASA

USLI will be followed for pressurized

air in vehicle, all containment systems

for the pressurized air will have high

safety factors and possible sources of

punctured minimized.

Launch

Operations

Explosive

motor

failure on

launch pad

Significant risk of

severe personnel harm

and fire

hazard. Significant risk

of loss of vehicle and

payload.

Defective or

damaged motor. 4 1 Follow MSDS storage

requirements. Insert and

prepare motor properly


initiated. Motor will be

inspected before launch.

Premature

motor

ignition (on

launch rail)

Significant risk of

personnel harm, fire

hazard, and

uncontrolled

launch. Potential loss

of vehicle and payload.

Improper storage or

handling of

motor. Improper

connection between

motor and igniter.

4 1 Maintain personnel a safe

distance away from

launchpad (300-500 ft

according to NAR High

Power Rocket Safety

Code). Follow MSDS

storage requirements. Insert

and prepare motor properly


initiated.

Motor fails

to ignite Vehicle will not launch. Improper connection

between motor and

igniter Motor igniter

falls out. Defective

motor. Motor is wet.

0 1 Follow NAR safety

procedure (60 second wait)

before accessing Launch

Vehicle following motor

failure to ignite. Check

connections prior to launch.

Insert and prepare motor

properly before launch

sequence is initiated. Motor

will be inspected before


launch.

Ground

Support

Equipment

Launch

Vehicle falls

off launch

rail

Launch Vehicle falls off

launch rail. If motor is lit,

Launch Vehicle will fire

off in uncontrolled

direction. Substantial risk

to personnel and

bystanders.

Launch rail

improperly secured

by RSO

4 1 RSO properly sets up

launch rail.

Premature

motor

ignition

(prior to

installation)

Significant risk of

personnel harm and

general fire hazard.

Heat source placed

near

motors. Improper

handling of motors.

4 1 Motors will be

provided the day of

launch. Will only be

handled by qualified

personnel. Follow

MSDS storage

requirements.

Premature

detonation of

ejection

charges (prior

to

installation)

Significant risk of

personnel harm and

general fire hazard.

Heat source placed

near

charges. Improper

handling of charges.

4 1 Charges will be

provided the day of

launch. Will only be

handled by qualified

personnel. Follow

MSDS storage

requirements.

Likelihood Definitions

Table 3.2.3 : Likelihood Definitions

Likelihood Definition Ranking

Remote Significant negligence and major defects required for hazard to occur. 1

Unlikely Significant negligence or defects required for hazard to occur. 2

Possible May occur despite proper safety measures and equipment checks taking place. 3

Likely Expected to occur despite proper safety measures and equipment checks taking

place. 4

Severity Definitions

Table 3.2.4 : Severity Definitions

Severity Definition Rank

Catastrophic Environment causes complete loss of system, or system causes significant permanent 4


damage to environment.

Major Major depletion of system functionality, substantial effect on environment. 3

Moderate Partial effect on system functionality, some effect on environment. 2

Minor Small effect on system functionality, mild environmental concern. 1

No Effect System stays intact, environmental conditions never altered. 0

The following table represents a list of environmental concerns associated with the launch vehicle.

Environmental Concerns

Table 3.2.5 : Environmental Concerns

Environmental

Concern Effects Causes Severi

ty Probab I l ity

Mitigation Verification

Motor

Chemicals Possible

contamination of

environment

surrounding Launch

Vehicle during

launch preparations

or launch sequence.

Improper handling

of

motor. Propellant

falls out of

dropped motor.

1 1 Transport and

load motor into

launch vehicle

properly,

according to

MSDS and

motor operator’s

manuals.

Have a

written

launch

procedure in

accordance

with MSDS

and motor

operator’s

manuals.

Impact of Motor

Ignition on

Launch Area

Possible fire or heat

damage to

immediate area

around launch

area. Potential fire

hazard.

No shield between

motor exhaust and

ground.

1 1 Plate mounted

on launch rail

between motor

to prevent motor

exhaust from

burning launch

area.

Verify launch

setup with

site

personnel.

Debris from

rapid unplanned

disassembly of

Launch Vehicle

Depending on the

scale of the vehicle

failure, debris

consistency and size

may vary. Debris

may consist of

carbon fiber and

“blue tube”

fragments.

Catastrophic

failure of

separation charges

or recovery

system of vehicle.

2 1 Launch Vehicle

systems will be

fully inspected

before launch to

mitigate

probability of

vehicle

failure. In the

event of a

failure, area will

be policed for

debris.

Checklist for

Launch

Vehicle

inspection.


Debris from

launch

preparation

Garbage and

disposable waste

(garbage bags,

wrappers, tape, etc.)

do not decompose

in a natural

environment and

must be collected.

Lack of proper

disposal plan for

consumables and

waste products.

2 2 Ensure any

waste is

properly

collected and

disposed of at

launch

site. Team area

will be policed

for debris and

waste, and any

found will be

properly

disposed of.

Use written

procedure for

proper waste

disposal to

ensure safety

during launch

site cleaning.

Precipitation Loss of electrical

system function due

to moisture. Motors

and separation

charges can also be

affected by

moisture, resulting

in a motor failure to

ignite, or a failure of

the Launch Vehicle

to deploy

parachutes.

Rain. 2 2 Team will

ensure vehicle is

protected from

precipitation

prior to

launch. Team

will avoid

launching when

rain is occurring

or predicted.

Launch

preparation

checklist will

include

ensuring a

safe forecast.

Bird Strike Risk of death or

serious injury to

bird. Risk of

significant damage

or complete loss of

Launch Vehicle,

depending on the

size of the bird hit.

Bird flies into path

of Launch Vehicle

on ascent.

4 1 Airspace above

launch area will

be

cleared. Team

will not launch

if there are

significant

numbers of birds

above the launch

site.

Launch

preparation

checklist will

include

checking for

birds.

High Winds Significant changes

to flight path or

vehicle stability due

to high

winds. Increased

wind drift after

parachute

deployment can

prevent recovery of

Launch Vehicle.

Inclement weather

or high winds. 2 2 Weather will be

monitored and

launches will

not take place if

wind speed is

too high.

Launch

preparation

checklist will

include

checking for

safe wind

speeds.


4. Payload Criteria

4.1. Design of Payload Equipment

4.1.1. Wheels The decision making process was completed through the use of comparison charts to weigh the pros and

cons of various options as seen in Table 4.1.1.1. The first design decision was to consider the number of

wheels on the rover. One option was for the rover to have four wheels. This idea has a number of pros

such as a robust, simple design, and the ability to maneuver over rough terrain. Some of the cons are that

it severely limits chassis size, limits wheel size, and the rover would be unable to self-right itself, which

would make the ejection system more complicated and dependent on payload orientation. A four-wheeled

rover would cause many severe problems, led the team to rule it out as one of the options for our final

design.

Another concept for wheel design on the rover is a tracked design, similar to tank treads. The pro that

come with this design is that it is best for rough terrain and turning. This rover would also have a robust

interface with the ejection system; another positive. Many of the limitations from the four-wheel design

carry are also applicable. It is a complex design and there is the possibility of the rover separating from

the tracks. The tracks would also limit the space and size of the rover because they take up space inside

the Launch Vehicle. The idea for a tracked design was also ruled out because of the complexity and the

challenges that face this design such as timing and idler wheels.

The third, and final, design is for a two-wheeled rover with one additional support appendage for

stabilization. This rover would be self-righting and allow a maximum chassis size compared to the

Launch Vehicle diameter. This design will allow for the rover to utilize counter-torque for self-righting

capability. The two-wheeled rover would allow for a larger wheel size and therefore utilize space in the

Launch Vehicle. Although this design has many positives, there are also some drawbacks. One of the

drawbacks is that the ejection process will be the amount of force between the wheel surface and launch

vehicle interface. This has the potential to cause problems on the wheel axle such as wheel deformation.

The material design will take into account the applied forces. Rough terrain is also more of a challenge

for the two-wheeled rover because it has less stability and ability to navigate the ground surface.

Although the design is less stable, a two-wheeled rover was determined to be the best design to complete

the challenge criteria.

Table 4.1.1.1: Decision Matrix for Wheel Design where each criterion was ranked and the design

alternative with the highest total value was considered the best design. Green indicates the chosen primary

design.

Criteria 4- Wheels 2-Wheels Treads

Size 2 3 1

Maneuverability 3 1 2

Stability 1 3 2

Complexity -2 -1 -3


Total 4 6 2

4.1.2. Ejection Another challenging design aspect is to determine how to eject the rover from the Launch Vehicle. Three

designs were evaluated in Table 4.1.2.1 one of the designs that was considered was a side hatch in the

Launch Vehicle which would be ejected by an airbag. This concept would bring the least amount of

potential damage to the rover during separation. A hatch also does not require separating the launch

vehicle sections on the ground. However, a hatch in the side of the rocket would severely reduce the

structural integrity of the Launch Vehicle. This hatch idea also depends on the Launch Vehicle landing in

the correct orientation so the hatch can open and the rover can properly exit the Vehicle. To add to the

negative aspect of this design, the rover may have trouble exiting the Launch Vehicle under its own

power. This is because the Launch Vehicle is cylindrical and creates a “U” shape when the hatch is

opened; thus creating a difficulty for the rover to exit. After considering all of these pros and cons, it was

determined that the hatch idea would be too difficult to execute successfully.

The second concept that was considered was a spring system to push the rover out of the Launch Vehicle.

The spring system would function independent of the orientation of the Launch Vehicle after it lands,

which is a pro compared to the hatch concept, which depends upon the orientation. The spring system

also occupies less space than the pneumatic piston design, which is explained below. It is desirable for the

ejection system to be as compact as possible so that more space can be allotted to the rover. A larger rover

will be more capable of completing its requirements. One of the pitfalls of the spring system is that it

cannot be gradually controlled, like the pneumatic piston can, so it has a greater chance of damaging the

rover. Another con is that the Motor Section needs to be ejected far enough away from the Payload

Section so that it does not obstruct the deployment of the rover. The spring system may not be able to

consistently and successfully push the rover out of the Launch Vehicle, and therefore was determined not

to be a viable option as an ejection method.

The final idea that was considered is the pneumatic piston. This is the concept that was ultimately

determined to be the most successful and the best idea to deploy the rover. The system does require space

for the pneumatic components, which is a downside. Also, there is a risk of puncture and the Motor

Section needs to be ejected far enough away from the Payload section as to not obstruct the rover exiting

the Launch Vehicle. With the pneumatic system, the rover can be deployed safely because the regulator,

valve, and speed controller can be controlled to adjust for the correct pressure and flow rate. This system

also does not depend on the orientation that the Launch Vehicle lands in. When considering all of these

factors and concepts, the best design to eject the rover from the Launch Vehicle is to use the pneumatic

piston.

Table 4.1.2.1: Decision Matrix for Rover Deployment, each criterion was ranked and design

alternative with least negative total value was the best design. Green indicates the chosen primary design.

Criteria Side Hatch Spring Forced Pneumatic Piston

Orientation 1 2 3

Structural Integrity of LV -3 -2 -1

Complexity -1 -2 -3


Space Required -1 -3 -2

Total -4 -5 -3

4.1.3. Solar Panel Deployment Once the rover is ejected from the Launch Vehicle, it will drive 5 feet and deploy foldable solar

panels. The decision matrix is displayed in Table 4.1.3.1. The first concept considered for solar panel

deployment is to have a spring-loaded system. A positive aspect of this design is that it would be simple

and compact. However, one risk is that the solar panels may be damaged when the spring-loaded

mechanism deploys due to the unregulated speed and no position feedback. This design is also considered

to be irreversible once the compressed springs are released. There is potential that the spring fails to open

and the system does not work.

An alternative design concept is a flexible solar panel. This design is relatively simple and can be bent in

many degrees of freedom, making it easy to fit inside the rocket and it is less likely to be damaged during

launch. Opposing the simplicity of this design, it is also incredibly bulky. The flexible solar panels would

be difficult to fit inside the Launch Vehicle; there is the possibility that they will get caught on the inside

of the Launch Vehicle during ejection. Another aspect that must be considered is the complexity of

mounting a rolled-up solar panel onto the rover.

The final design for the solar panel ejection method is a servo motor-driven fan array deployment. This

design takes up more horizontal space in the chassis and it is more complicated compared to other designs

and therefore it is more prone to failure. The motor-driven deployment is also the most compact in the

vertical direction and has the most controlled and regulated motion. The motor selected will be designed

to function both forward and reverse such that the solar panels can be folded back to starting position for

ease of testing. It has the highest probability to succeed of the three systems, and therefore is the system

that was chosen to deploy the solar panels on the rover.

Table 4.1.3.1: Decision Matrix for Solar Panel Deployment where each criterion was ranked and the

design alternative with the highest total value was considered the best design. Green indicates the chosen

primary design.

Criteria Spring-Loaded Flexible PV Servo Fan

Complexity -1 -2 -3

Space 2 1 3

Control 1 2 3

Total 2 1 3

4.2. System Level Design Review 4.2.1. Electronics

The barometer is connected to analog inputs on the Arduino Nano to record the altitude. The data is then

transmitted via the XBEE radio device which is connected to the microcontroller. It is transmitted to the

ground station, which will then be parsed into readable values, and will also transmit the current

activation state to the team. Once the rover is ejected using the pneumatic piston, the accelerometer and


gyroscope is used to measure and record the position and the chassis orientation of the rover starting from

its initial point.

4.2.2. Solar Panel

A flexible 6V 1W solar panel will be deployed using a servo motor driven fan array to rotate the panels

from 0 to 180 degrees. The panels will be connected to the onboard electronics, and the energy generated

will be used to light up an indicator light. The motor is designed to function both forward and backwards

in order for the solar panels to be folded back into its initial position.

4.2.2. Solar Panel A flexible 6V 1W solar panel will be deployed using a servo motor driven fan array to rotate the panels

from 0 to 180 degrees. The panels will be connected to the onboard electronics, and the energy generated

will be used to light up an indicator light. The motor is designed to function both forward and backwards

in order for the solar panels to be folded back into its initial position.

4.2.3. Wheels

The rover has 2 wheels that have a diameter of 5.9 inches each. The cores of the wheels are 3D printed

and then sprayed with a rubber coating to provide traction on the rough terrain. Two Vex 2 Wire Motors

are used to drive the wheels of the rover.

4.2.3. Payload Ejection System

Prior to launch, a Peregrine CO2 ejection system pressurizes the pneumatic system. A compact regulator

will drop the pressure to approximately 35 psi. A 3 ported, 2 way internally piloted normally closed valve

following the regulator will remain shut until a 12V signal is received from the PES electronics bay (after

a signal from the ground station is received). A speed controller (set prior to launch) will control the flow

of the CO2 once the valve is opened. Bellows, constructed from polyester film, will be expanded, pushing

the forward bulkhead and the rover from the rocket. The bellows will be clamped to the forward and rear

bulkheads. The forward bulkhead has three bushings epoxied to it. These bushings slide along steel rods

which are affixed to the rear bulkhead. All components are connected with nylon tubing (1/4” OD and

11/64” ID) and barbed fittings (¼ NPT).

4.3. Drawings and Specifications for Components and Assembly The Payload Ejection System has two states: retracted and extended. The major dimensions of both states

are shown below in Figure 4.3.1. Note that the polyester film is not shown.


Figure 4.3.1: Extended and Retracted Payload Ejection System

The PES contains 3 bulkheads: The forward bulkhead, the rear bulkhead, and the pneumatic base

bulkhead. The forward and rear bulkheads have the same profile and are shown below in Figure 4.3.2.

Note that the rear bulkhead has no bushings, but rather epoxy to retain the steel rods.

Figure 4.3.2: Profile of forward and rear bulkheads


The pneumatic components are mounted to the pneumatic base bulkhead and are shown below in Figure

4.3.3.

Figure 4.3.3: Pneumatic layout

The electronics to control the pneumatic system and to communicate with the ground station and the rover

are shown below in Figure 4.3.4. They are mounted to the pneumatic components bulkhead.

Figure 4.3.4: Electronic layouts

An engineering drawing of the rover body is shown below in Figure 4.3.5.


Figure 4.3.5: Drawing of Rover Body

The profile of the wheel has the following major dimension as shown in Figure 4.3.6. Solenoid latches are

shown.

Figure 4.3.6: Profile of the wheel


The length of the rover is shown below in Figure 4.3.7.

Figure 4.3.7: Length of the Rover

4.4. Payload Component Interaction Description

4.4.1. Payload Component Interaction Overview

The NU-Frontiers payload consists of two major sections which must interface with one another. The

payload section includes the NU-Frontiers Rover (NUFR) and Payload Ejection System (PES). These

sections must effectively communicate with one another to assure successful payload ejection and on

ground performance.

To begin, we will start with a high level overview of the NUFR payload, what NUFR consists of, and

how these components are composed to create a functioning rover.

4.4.2. NUFR Component Interaction


2x 7.4V 1000mAh LiPo

GY-521 Accelerometer and Gyroscope

Adafruit Ultimate GPS Breakout

KY-019 Relay

Sunfounder PCA9685 Servo Driver

Arduino Nano

XBEE PRO 60mW

Flexible 6V 1W Solar Panel

DS1307 Real Time Clock (RTC)

IG-406D-2341 Key Lock Switch

PES Electronic Components Arduino Nano V3


IG-406D-2341 Key Lock Switch

KY-019 Relay

XBEE Pro S3B


S3170G Servo (Tail Servo)

BLS-173SV Servo (Solar Servo)

Vex 2 Wire 393 Motor (Drive Motor)

BMP180 Barometer/Altimeter

Relieving air regulator

3 ported, 2 way internally piloted solenoid

valve

The NU-Frontiers Rover is designed with autonomous function in mind. Once the rover exits the launch

vehicle and separates from the PES ejection system, it must complete tasks without outside interference.

Keeping that in mind, the rover must still be able to interface with the PES ejection system, and by proxy,

the ground station.

This system consists of a microcontroller, keylock switch, radio device, relay, gyroscope, GPS, servo

driver, and RTC. During launch the system will remain in a ‘sleep state’, waiting for the signal to activate,

while in this sleep state the GPS position will be relayed to the ejection system back to the ground station.

On signal reception by the ejection system, another signal will be sent to the rover via the ejection system,

activating the system.


On activation, the RTC will begin counting. During ejection tests a safe time estimate for ejection will be

determined, and the Arduino - RTC system will wait until that time is reached. Once the timer has

reached this predetermined count, the rover will begin operation. First the servo driver will flip out a ski to provide counter moment for rover movement, following this

event the relay onboard the rover will activate, connecting the wheel motors to ~7.4V of power (two 7.4V

LiPos connected in parallel). At this point the RTC will begin another count, this time to a predetermined

estimate (based on tests for rover distance), after this time has passed, the relay will switch to its ‘off’

position, cutting off wheel motor power. Next the PWM driver will actuate the solar servo, deploying

solar panels and completing the rover task.

4.4.3. PES Component Interaction This system consists of an Arduino Nano microcontroller, XBee Pro 60mW, KY-019 relay, two 7.4 V

1000mAh LiPo Batteries, BMP180 barometer, 5V and 12V voltage regulators, and a keylock switch.

These components are all soldered to a protoboard. The pneumatic components consist of the CO2

cartridge (standard 12g size), a Peregrine CO2 Ejection System, a compact pressure regulator with an

internal relief valve, a 3 ported, 2 way internally piloted solenoid valve, a speed controller, and a

polyester bellow. Structurally, the system consists of 3 bulkheads and 3 steel rods. The bulkheads are the

pneumatic base bulkhead will be set screwed in place. The rear bulkhead (between the front bulkhead and

the pneumatic base bulkhead) will be epoxied in place. The forward bulkhead will not be fixed to the

body tube; instead it will be free to slide along the steel rods affixed to the rear bulkhead. All bulkheads

will be within the body and will have a diameter slightly smaller than the body tube ID (~5.95 inches). Before launch, the keylock switch (wired in series with the battery and microcontroller) will be turned,

powering the system. At this point the barometer, connected to analog inputs on the microcontroller will

begin recording altitude. This data will be transmitted via XBEE radio device connected to the

microcontroller. In turn, the ground station will receive this data and parse it as human readable values.

Additionally the radio will transmit the current ‘activation state’ (see definition below) for the rover. The activation state will be set as 0 from rocket launch to landing, once the activation signal is sent from

the ground station and received by the PES; the activation state will be set as 1. On system ‘activation’, a signal will be sent from the high powered XBEE in PES to the low powered

XBEE on the rover. On signal reception, the ejection system will activate. On activation the relay will be

powered, which will in turn send a 12V signal to the solenoid valve, starting the ejection process. Prior to vehicle assembly, a CO2 ejection system will pressurize the pneumatic system. After receiving a

signal from the ground station and changing the state to “activation state” to 1, the valve will open and fill

the bellows. The pressure within the bellows will push the rover out of the launch vehicle. Once rover ejection has completed the PES will work as a proxy, relaying data from the rover to the

ground station.

4.5. Payload Integration Plan The full scale rocket contains a payload section that is placed into the body tube of the launch vehicle.

This payload section contains the NU Frontiers rover (NUFR) and the Payload Ejection System (PES).

Two bulkheads of the PES are rigidly connected to the body tube of the launch vehicle by two bulkheads.

The pneumatic base bulkhead is set screwed in place and the rear bulkhead is epoxied in place. The


NUFR is connected radially to the PES by steel rods. It is axially constrained in one direction by the rear

bulkhead and by the coupler tube in the other (this coupler is shear screwed to the Lower Avionics

Section prior to separation). The coupler tube is prevented from falling out of the body tube prematurely

by solenoid pull latches. The solenoid latch shafts are in slots in the coupler and body tube in order to

protect the shafts from the shock of the black powder separation charge. The shock from the separation

will instead be transferred to set screws position immediately fore of the coupler. The solenoids will pull

in when they receive a signal from the ground station. The NUFR and PES are integrated via a wireless radio link using two XBee radios the radio in the PES

will also be used to communicate with the ground station. This payload section is 18’’ long and 6’’ in

diameter which fits into the body tube of the rocket. The connections between the NUFR and the launch

vehicle include the bulkhead that anchors the aft end of the PES, and the bulkhead that protects the NUFR

and PES from the separation charges in the upper section of the vehicle. The NUFR contains a PCB that

contains the electronics that measures the flight data and sends it to the ground station. A Peregrine CO2

ejection system will be placed in the body tube to pressurize the pneumatic system, which will then eject

the rover upon landing. The NUFR is ejected from the rocket via a pneumatic bellows piston that pushes

the payload out of the rocket body and then NUFR performs its operations on the ground.

4.6. Demonstration of Payload Design Completion The payload completion depends on various factors: final decisions on sensors/motors/servos/wheels,

payload material, circuit board design, mission plan, and ejection plan. These factors will be described in

the following plan.

Figure 4.6.1: Rendering of Rover, Solar Panels Retracted


Figure 4.6.2: Rendering of Rover, Solar Panels Deployed

Sensors A full sensor suite has been decided on by the Payload Electronics Team, the suite is comprised of:

Arduino Nano V3

7.4V 1000mAh Floureon LiPo Battery

2.4GHz Duck Antennae

CR1220 Battery

Keylock Switch

Arduino Ultimate GPS Breakout

Accelerometer MPU6050

Keyes Relay

Sunfounder Servo Driver

Motors/Servos The motors and servos to be used in the rover design are listed below with the respective components

which they drive Two Vex 2 Wire Motors will directly drive the wheels of the rover

The counter torque will be deployed and held in place by a Futaba S3170G Dig Retract Servo

The solar panels will be deployed by a Futaba BLS-173SV Servo

Wheels After careful deliberation regarding the terrain of the launch site, durability, and weight/size constrictions

the NU Frontiers team decided on using custom 3D printed wheels with a sprayable rubber coating to

provide additional traction on rough surfaces. Circuit Board Design


As highlighted in the drawings section of the CDR, the PCB design for the NU Frontiers rover has been

completed and is in its final form. The PCB will allow for maximization of space, optimization of

soldering connections, and will reduce the chance of solder bridges which would be a failure mode if the

team used a proto board instead Mission Plan The mission plan for the NU Frontiers Rover has been clearly defined in the CDR. Ejection Plan The ejection plan for the NU Frontiers Rover, utilizing the PES ejection system has been clearly defined

in the CDR.

4.7. Payload Electrical Drawings and Dimensions Included in this section is the Fritzing diagram and PCB schematic for the NU-Frontiers Rover.

Figure 4.7.1: Fritzing diagram for NU-Frontiers Rover

Not seen above:

2x 7.4v LiPo connected in parallel to provide a redundant power source

1x Keylock switch connected in series to circuit, to be armed on pad

The above schematic is indicative of the final PCB design; all hookups will be identical for the final

design. Switches and wires will be separate from this design, mounted on the rover itself and wired into

the PCB.


The Fritzing software was used to design the payload layout. As Fritzing is a relatively new software,

there were some hurdles to overcome. Especially the DRC (Design Rule Checks) for the PCB design

were carefully check to make sure that no wires crossed or created shorts.

After several attempts, some utilizing the Autoroute functionality provided with the Fritzing software, the

team determined that it would best to manually route all connections.

Screw holes were placed on the outside of the PCB; it is the team’s intention to mount the PCB directly

within a side of the rover.

Figure 4.7.2: Initial PCB layout, based off autoroute functionality within the Fritzing software


Figure 4.7.3: Final PCB layout, manually routed with added via connections

4.8. Payload Block Diagrams The diagram below describes the connections between the different components used in the PES, NUFR

and ground station and the connections that are made between these components.


Figure 4.8.1: Block Diagram of Payload and Ground Station

The diagram below shows the program that executes on the Arduino that controls the NUFR

Figure 4.8.2: Arduino Program Diagram


4.9. Payload Battery and Power Consumption The NU Frontiers Rover (NUFR) will be powered by two 7.4V 1000 mAh li-po batteries connected in

parallel to increase the current for use by the motors powering the drive system. This gives a total of 2000

mAh of total power for the rover. The Payload Ejection System will also be powered by an additional two

identical 7.4V 100mAh Li-Po batteries.

Figure 4.9.1: 7.4V 1000mAh Li-Po

Table 4.9.1: Device Listing and Power Consumptions

Device Power Consumption

Xbee 1W

Arduino Nano 0.6W

GPS 0.1 W

Each battery is capable of supplying 7.4 Watt hours of power, with 2 batteries for a total power of 14.8

Watt hours. This means that the electronics in the PES can operate for approximately 6 hours before

power becomes a concern for the payload deployment phase of the flight.

4.10. Switch and Indicator Wattage and Location In order to arm the rover, a keylock switch will be used. The rover will be armed on the pad via a hole in

the body tube of the launch vehicle.


Figure 4.10.1: A Keylock Switch

The team will be notified of successful arming by a tone pattern played via the piezo buzzer on the rover

itself. On successful arming, the buzzer will emit the tone: 110110110000-110110110000-110110110000-110110110000-110110110000-... Where 1 is on, 0 is off, each digit represents half a second, and the pattern is played on repeat. Different tones will eventually replace the ‘power tone’ to report basic data back to the team on rover

retrieval. The keylock switch will be located on the rear of the rover, while the batteries will be housed on the left

side of the rover.


Figure 4.10.2: 2x LiPo battery housing displayed in gold

LiPo batteries will be hooked in parallel or series depending on the voltage required to functionally power

the wheel motors.

4.11. Payload Justification Many of our payload’s attributes that contribute to its uniqueness are original designs chosen in order to

accommodate the functionality of the rover itself. Examples of aspects of our experiment that we feel are

unique include our 2-wheeled design, dynamic support strut, home-made wheels, and pneumatic ejection

system. The rover is 2-wheeled because a 4-wheel or 3-wheel rover would have serious size constraint being

stored inside the launch vehicle. By reducing the number of wheels to 2, the rover can fit horizontally

within the payload bay, and the wheels can therefore approach the inner diameter of the air frame. This is

necessary for adequate traversal of the hummocky terrain. The dynamic support strut is a component the need for which arises from the fact that the 2-wheeled rover

would otherwise be unable to counteract the torque applied to its wheels by the ground. If the structure of

the rover were so limited in the amount of torque it could provide, the chassis would spin with respect to

the wheels, rather than the wheels rotating with respect to the ground. The strut therefore provides the

necessary torque to control the wheels, while also being storable to reduce the rover’s profile. The wheels will be 3D printed because our ejection system requires several holes to be present in the

wheels. It will also be cheaper and easier to prototype 3D printed wheels. The wheels will be sprayed with

rubber to make them water-proof, while also providing an amount of traction comparable to commercially

available wheels, all the while maintaining the specific dimensions that our ejection system requires. Finally our pneumatic ejection system is fairly unique as it utilizes a polyester bellows that will be

inflated. This is particularly useful when space is constrained in the axial direction of the rocket, because


most other ejection methods require as much physical space as the rover in order to fully deploy it. A

pneumatic system, on the other hand, is compressible, possible to 25% of its expanded length.


5. Project Plan

5.1. Testing

5.1.1. Payload Tests

To verify the functionality of the payload we will need to perform tests that include radio range

and data tests; signal testing on the rover and payload ejection system; verifying the code

functionality on the rover, payload ejection system, and ground station; running the payload

components at both full power and idle power to find the battery life in both of these scenarios;

testing the ability for the payload to withstand the forces at separation, drogue deploy, main chute

deploy, and landing; the ability for the rover to drive on rough terrain; finally, the ability of the

solar panels to deploy.

5.1.2. Rover Test Plans

Radio Communications Testing

A ground station and basic payload test were constructed with the following parameters

o Ground Station

USB Connection to Laptop computer

XBEE Pro 60mW

o Payload

Arduino Nano V3

XBEE Pro 60mW

7.4V 1000mAh LiPo

o During this test the ground station was placed at a set spot with linear sight on the

payload, packets were then transmitted from the payload device to the ground

station with various distances (NOTE this was done in an urban environment,

buildings impacted signal strength) signal reception tests were as follows

10m - Signal Reception






200m - Spotty Signal Reception (1 packets per ~1s)

250m - Spotty Signal Reception (1 packet per ~3s)

300m - Little to no Signal Reception (1 packet per ~20+ s)

A basic “dummy” payload was loaded into the subscale launch vehicle and a packet

transmit/reception test was conducted with the following parameters using a ground

station


o Ground Station


XBEE Pro 60mW

o Payload

XBEE pro 60mW

7.4V 1000mAh LiPo

BMP180 Altimeter/Barometer Sensor

Arduino Nano V3

Accelerometer

o During this test the NU Frontiers sub-scale rocket was launched to an altitude of

2400 ft, during launch the “dummy” payload attempted to transmit altitude, and

accelerometer values. Transmission was completely lost at 2100 ft, with packets

being dropped starting around 1000 ft. The team decided to implement a higher

powered XBEE for the final design.

o Additionally, a signal transmission from the ground station was tested,

unfortunately this failed repeatedly until about 100 yards from the rover, where

signal was found and the transmission was received.

Battery Duration Testing We plan to test the duration that the batteries, in both the payload ejection system and the

rover, will last when under full load and idle pre-launch load.

o Ground Station


XBEE Pro S2C

o Payload Ejection System (PES)

Arduino Nano V3

XBEE Pro S2C


BMP180 Barometer

Valve

o NU Frontiers Rover (NUFR)

Arduino Nano V3

XBEE Pro S2C


Accelerometer/Gyroscope

2x Vex Motors

2x Servos

Real Time Clock

GPS module

Servo Driver

o During this test we will run the system at idle and at full power draw to see how

long the systems are able to operate. This test will be considered a success if we

are able to determine the duration that the payload can operate on a fully charged

set of batteries.

o It is necessary to know this information on the battery duration to determine if we

will need to modify the payload to function after the required minimum of 1 hour

on the launch pad, with safety margin and power for deployment and roving.


5.2. Requirements Compliance

Verification Plan

Table 5.2.1 : Verification Plan

Number NASA Derived Requirement Requirement Completion

Plan Verification

1.1 Students on the team will do 100% of the

project, including design, construction, written

reports, presentations, and flight preparation

with the exception of assembling the motors

and handling black powder or any variant of

ejection charges, or preparing and installing

electric matches (to be done by the team’s

mentor).

Northeastern University’s

Chapter of AIAA takes pride in

exposing students to hands-on

engagement. Students in the

NASA USLI Competition will

have taken initiative and

leadership in the execution of

this project with the close

supervision of our mentor.

Self-Evaluation of

student involvement at

milestones

1.2 The team will provide and maintain a project

plan to include, but not limited to the following

items: project milestones, budget and

community support, checklists, personnel

assigned, educational engagement events, and

risks and mitigations.

The Northeastern University

Team has several members

committed to creating and

monitoring plans, a schedule,

and a budget to ensure these

requirements are met.

Periodic review and

evaluation of progress

relating to

aforementioned

requirements

1.3 Foreign National (FN) team members must be

identified by the Preliminary Design Review

(PDR) and may or may not have access to

certain activities during launch week due to

security restrictions. In addition, FN’s may be

separated from their team during these

activities.

There are no FN team members

on the Northeastern University

team.

Verified

1.4 The team must identify all team members

attending launch week activities by the Critical

Design Review (CDR). Team members will

include:

The team members attending

launch week activities have been

determined and notified.

N/A

1.4.1. Students actively engaged in the project

throughout the entire year. Will be verified by attendance

record, engagement and

contribution.

N/A

1.4.2. One mentor (see requirement 1.14). Identified as Robert DeHate. Verified

1.4.3. No more than two adult educators. N/A N/A

1.5 The team will engage a minimum of 200

participants in educational, hands-on science,

technology, engineering, and mathematics

(STEM) activities, as defined in the

Educational Engagement Activity Report, by

FRR. An educational engagement activity

report will be completed and submitted within

two weeks after completion of an event. A

sample of the educational engagement activity

report can be found on page 31 of the

handbook. To satisfy this requirement, all

events must occur between project acceptance

The team has a dedicated STEM

outreach coordinator who

organizes STEM education

events in conjunction with

Northeastern University’s Center

for STEM education such as

field trips for local high schools,

science fairs, and paper rockets

demonstrations. This STEM

coordinator will also be

responsible for submitting

engagement activity reports.

Engagement will be

measured by attendance

by K-12 students and

counted only if 2 or more

team members attend the

event.


and the FRR due date.

1.6 The team will develop and host a Web site for

project documentation. The website is hosted at

neu.edu/aiaa. A team member is

dedicated as webmaster and will

upload documents and maintain

the site as necessary.

N/A

1.7 Teams will post, and make available for

download, the required deliverables to the team

Web site by the due dates specified in the

project timeline.

The webmaster will upload

documents as necessary. N/A

1.8 All deliverables must be in PDF format. The team lead will ensure that all

deliverables are in PDF format

before submitting.

Visual

1.9 In every report, teams will provide a table of

contents including major sections and their

respective Sub-sections.

The team lead will ensure that all

reports include a table of

contents.

Visual

1.10 In every report, the team will include the page

number at the bottom of the page. The team lead will ensure all

reports include page numbers. Visual

1.11 The team will provide any computer equipment

necessary to perform a video teleconference

with the review panel. This includes, but is not

limited to, a computer system, video camera,

speaker telephone, and a broadband Internet

connection. Cellular phones can be used for

speakerphone capability only as a last resort.

The team lead will ensure that

the proper equipment is procured

and available well in advance of

the teleconference. A meeting

space will be booked with the

Department of Mechanical and

Industrial Engineering at

Northeastern University.

N/A

1.12 All teams will be required to use the launch

pads provided by Student Launch’s launch

service provider. No custom pads will be

permitted on the launch field. Launch services

will have 8 ft. 1010 rails, and 8 and 12 ft. 1515

rails available for use.

Rail buttons for 1010 rail will

be used. N/A

1.13 Teams must implement the Architectural and

Transportation Barriers Compliance Board

Electronic and Information Technology (EIT)

Accessibility Standards (36 CFR Part 1194)

Subpart B-Technical Standards

(http://www.section508.gov): § 1194.21

Software applications and operating systems. §

1194.22 Web-based intranet and Internet

information and applications.

1.14 Each team must identify a “mentor.” A mentor

is defined as an adult who is included as a team

member, who will be supporting the team (or

multiple teams) throughout the project year,

and may or may not be affiliated with the

school, institution, or organization. The mentor

must maintain a current certification, and be in

good standing, through the National

Association of Rocketry (NAR) or Tripoli

Rocketry Association (TRA) for the motor

Robert DeHate has been

identified as the mentor for our

team.

Verified

http://www.section508.gov/


impulse of the launch vehicle and must have

flown and successfully recovered (using

electronic, staged recovery) a minimum of 2

flights in this or a higher impulse class, prior to

PDR. The mentor is designated as the

individual owner of the rocket for liability

purposes and must travel with the team to

launch week. One travel stipend will be

provided per mentor regardless of the number

of teams he or she supports. The stipend will

only be provided if the team passes FRR and

the team and mentor attends launch week in

April.

2.1 The vehicle will deliver the payload to an

apogee altitude of 5,280 feet above ground

level (AGL).

OpenRocket was used to predict

the apogee of the vehicle as

5310 ft and the design will take

into account this constraint.

Appropriate simulation

conditions will be used and

parameters such as weight and

center of mass will be double-

checked against real

measurements taken in the lab.

Avionics bays will

include redundant

altimetry data to verify

apogee at each launch.

2.2 The vehicle will carry one commercially

available, barometric altimeter for recording the

official altitude used in determining the altitude

award winner. Teams will receive the

maximum number of altitude points (5,280) if

the official scoring altimeter reads a value of

exactly 5280 feet AGL. The team will lose one

point for every foot above or below the required

altitude.

The team will attempt to

achieve an apogee of 5280 with

a combination of design, testing

by launch and iteration to

produce a robust final vehicle

before FRR.

The vehicle will carry

the competition

altimeter.

2.3 Each altimeter will be armed by a dedicated

arming switch that is accessible from the

exterior of the rocket airframe when the rocket

is in the launch configuration on the launch

pad.

A rotary switch that is

accessible from the exterior will

be used to arm the competition

altimeter as well as all other

altimeters.

N/A

2.4 Each altimeter will have a dedicated power

supply. A 9V battery will be installed in

every avionics bay that is

dedicated only to the altimeter.

N/A

2.5 Each arming switch will be capable of being

locked in the ON position for launch (i.e.

cannot be disarmed due to flight forces).

The rotary switch will be held

in the ON position and is not

capable of being disarmed in-

flight.

N/A

2.6 The launch vehicle will be designed to be

recoverable and reusable. Reusable is defined

as being able to launch again on the same day

without repairs or modifications.

The recovery system will be

designed to slow the vehicle’s

descent to protect it from

damage due to impact with the

surface. No subsystem is altered

during launch and can therefore

be recovered and reused.

N/A

2.7 The launch vehicle will have a maximum of The Launch Vehicle will have N/A


four (4) independent sections. An independent

section is defined as a section that is either

tethered to the main vehicle or is recovered

separately from the main vehicle using its own

parachute.

two independent sections. The

Nose Cone section will fall with

a main parachute independently.

The Booster, Payload, and

Avionics Bay Sections will all

fall tethered together with a

main and drogue parachute.

2.8 The launch vehicle will be limited to a single

stage. The Launch Vehicle will have a

singular motor stage. Visual

2.9 The launch vehicle will be capable of being

prepared for flight at the launch site within 3

hours of the time the Federal Aviation

Administration flight waiver opens.

The team will prepare the

Launch Vehicle and work

together to get the rocket

prepared in under 3 hours.

N/A

2.10 The launch vehicle will be capable of

remaining in launch-ready configuration at the

pad for a minimum of 1 hour without losing the

functionality of any critical on-board

components.

The electronics and systems on

the Launch Vehicle and Payload

will not be altered by the heat or

other environmental impacts

caused by waiting on the pad.

Testing and

demonstration: Full

scale test flight

2.11 The launch vehicle will be capable of being

launched by a standard 12-volt direct current

firing system. The firing system will be

provided by the NASA-designated Range

Services Provider.

The team will do test launches

to ensure that the launch vehicle

is able to launch with a standard

12-volt direct current firing

system.

Testing and

demonstration: Full

scale test flight

2.12 The launch vehicle will require no external

circuitry or special ground support equipment

to initiate launch (other than what is provided

by Range Services).

During the design and build

process the team will ensure

that the launch vehicle is

independently initiated.

Demonstration: Full

scale test flight

2.13 The launch vehicle will use a commercially

available solid motor propulsion system using

ammonium perchlorate composite propellant

(APCP) which is approved and certified by the

National Association of Rocketry (NAR),

Tripoli Rocketry Association (TRA), and/or the

Canadian Association of Rocketry (CAR).

The team implementing a CTI

motor which is commercially

available, uses APCP, and is

approved by specified rocketry

organizations.

N/A

2.13.1. Final motor choices must be made by

the Critical Design Review (CDR). The team has chosen the

Cessaroni L1115. N/A

2.13.2. Any motor changes after CDR must be

approved by the NASA Range Safety Officer

(RSO), and will only be approved if the change

is for the sole purpose of increasing the safety

Margin.

The team will notify NASA if

they plan on changing the motor

after CDR.

N/A

2.14 Pressure vessels on the vehicle will be

approved by the RSO and will meet the

following criteria:

The team will communicate all

information regarding the

pressure vessel to RSO.

Demonstration

2.14.1. The minimum factor of safety (Burst or

Ultimate pressure versus Max Expected

Operating Pressure) will be 4:1 with supporting

design documentation included in all milestone

reviews.

Calculations and documentation

will be reviewed and presented

to NASA at each milestone.

Analysis

2.14.2. Each pressure vessel will include a A pressure relief valve meeting Analysis


pressure relief valve that sees the full pressure

of the valve that is capable of withstanding the

maximum pressure and flow rate of the tank.

said requirements will be

identified and used in the design

of the system.

2.14.3. Full pedigree of the tank will be

described, including the application for which

the tank was designed, and the history of the

tank, including the number of pressure cycles

put on the tank, by whom, and when.

The team will ascertain and

present this information to

NASA.

N/A

2.15 The total impulse provided by a College and/or

University launch vehicle will not exceed 5,120

Newton-seconds (L-class).

L1395-BS made by CTI, where

the specifications say that it will

not exceed 5,120 Newton-

seconds.

The specifications

provided by the

manufacturer will verify.

2.16 The launch vehicle will have a minimum static

stability margin of 2.0 at the point of rail exit.

Rail exit is defined at the point where the

forward rail button loses contact with the rail.

The static stability margin was

calculated using OpenRocket. A

stability of 3.3 is given.

Analysis

2.17 The launch vehicle will accelerate to a

minimum velocity of 52 fps at rail exit. The velocity at rail exit was

calculated using OpenRocket.

69.9 ft/s is given.

Analysis

2.18 All teams will successfully launch and recover

a subscale model of their rocket prior to CDR.

Subscales are not required to be high power

rockets.

The original subscale rocket

launch was not considered

successful. Therefore a second

subscale was launched and was

successful.

Demonstration

2.18.1. The subscale model should resemble

and perform as similarly as possible to the full-

scale model; however, the full-scale will not be

used as the subscale model.

The subscale model was

constructed such that it would

perform similarly to the full

scale model.

Demonstration

2.18.2. The subscale model will carry an

altimeter capable of reporting the model’s

apogee altitude.

The subscale contained

PerfectFlite Stratologgers to

record the model’s apogee

altitude.

Demonstration

2.19 All teams will successfully launch and recover

their full-scale rocket prior to FRR in its final

flight configuration. The rocket flown at FRR

must be the same rocket to be flown on launch

day. The purpose of the full-scale

demonstration flight is to demonstrate the

launch vehicle’s stability, structural integrity,

recovery systems, and the team’s ability to

prepare the launch vehicle for flight. A

successful flight is defined as a launch in which

all hardware is functioning properly (i.e. drogue

chute at apogee, main chute at a lower altitude,

functioning tracking devices, etc.). The

following criteria must be met during the full-

scale demonstration flight:

The team will construct and

launch a full scale launch

vehicle prior to FRR. This

launch will demonstrate all

required criteria.

Demonstration

2.19.1. The vehicle and recovery system will

have functioned as designed. The flight events and launch

vehicle will be inspected to

check that it has functioned as

designed.

Inspection


2.19.2. The payload does not have to be flown

during the full-scale test flight. The following

requirements still apply:

N/A N/A

2.19.2.1. If the payload is not flown, mass

simulators will be used to simulate the payload

mass.

A simulated mass equal to the

mass of the payload, 2.8 kg,

will be used in test launches if

the payload is not flown.

N/A

2.19.2.1.1. The mass simulators will be located

in the same approximate location on the rocket

as the missing payload mass.

In the case a simulated mass is

used in place of the payload, the

mass will be in the approximate

location of the missing payload.

N/A

2.19.3. If the payload changes the external

surfaces of the rocket (such as with camera

housings or external probes) or manages the

total energy of the vehicle, those systems will

be active during the full-scale demonstration

flight.

External surface changes and

total energy management will

be implemented during the full-

scale demonstration flight if

applicable.

N/A

2.19.4. The full-scale motor does not have to be

flown during the full-scale test flight. However,

it is recommended that the full-scale motor be

used to demonstrate full flight readiness and

altitude verification. If the full-scale motor is

not flown during the full-scale flight, it is

desired that the motor simulates, as closely as

possible, the predicted maximum velocity and

maximum acceleration of the launch day flight.

The full scale motor or a similar

motor will be used in the full

flight readiness test to simulate

the predicted flight data of the

launch day flight.

N/A

2.19.5. The vehicle must be flown in its fully

ballasted configuration during the full-scale test

flight. Fully ballasted refers to the same amount

of ballast that will be flown during the launch

day flight. Additional ballast may not be added

without a re-flight of the full-scale launch

vehicle.

The vehicle will be flown in its

fully ballasted configuration

during the full-scale test flight.

If additional ballast is desired,

another full-scale test will be

flown with the additional ballast

added

N/A

2.19.6. After successfully completing the full-

scale demonstration flight, the launch vehicle or

any of its components will not be modified

without the concurrence of the NASA Range

Safety Officer (RSO).

No changes will be made to the

launch vehicle or its

componentes after a successful

full-scale flight without the

concurrence of the NASA

Range Safety Officer

N/A

2.19.7. Full scale flights must be completed by

the start of FRRs (March 6th, 2018). If the

Student Launch office determines that a re-

flight is necessary, then an extension to March

28th, 2018 will be granted. This extension is

only valid for re-flights; not first-time flights.

A full scale flight will be

completed by March 6th, 2018.

If required by NASA, a re-flight

will be flown by March 28th,

2018

N/A

2.20 Any structural protuberance on the rocket will

be located aft of the burnout center of gravity. Any structural protuberance on

the rocket will be located aft of

the burnout center of gravity.

Analysis/Inspection

2.21 Vehicle Prohibitions

2.21.1. The launch vehicle will not utilize

forward canards. The team will not put forward

canards on the launch vehicle. Visual



forward firing motors. The team will not put forward

firing motors on the launch

vehicle.

Visual


motors that expel titanium sponges (Sparky,

Skidmark, MetalStorm, etc.

The team will not use a motor

that expels titanium. Visual


hybrid motors. The team will not use a hybrid

motor on the Launch Vehicle. Visual

2.21.5. The launch vehicle will not utilize a

cluster of motors. There will not be a cluster of

motors on the Launch Vehicle Visual


friction fitting for motors. There will not be friction fitting

for motors on the Launch

Vehicle.

Visual

2.21.7. The launch vehicle will not exceed

Mach 1 at any point during flight. The flight will be simulated

using OpenRocket to ensure

that the launch vehicle does not

exceed mach 1

Analysis

2.21.8. Vehicle ballast will not exceed 10% of

the total weight of the rocket. Vehicle ballast will not exceed

10% of the total weight of the

rocket

Inspection

3.1 The launch vehicle will stage the deployment of

its recovery devices, where a drogue parachute

is deployed at apogee and a main parachute is

deployed at a lower altitude. Tumble or

streamer recovery from apogee to main

parachute deployment is also permissible,

provided that kinetic energy during drogue-

stage descent is reasonable, as deemed by the

RSO.

The launch vehicle will stage

deployment of its recovery

devices, deploying drogue

parachutes at apogee and larger

main parachutes at a lower

altitude. Kinetic energy will be

approved by the RSO.

Demonstration

3.2 Each team must perform a successful ground

ejection test for both the drogue and main

parachutes. This must be done prior to the

initial subscale and full-scale launches.

Successful ground ejection tests

will be performed for all

recovery systems prior to

subscale and full-scale

launches.

Demonstration

3.3 At landing, each independent sections of the

launch vehicle will have a maximum kinetic

energy of 75 ft-lbf.

OpenRocket and hand

calculations will predict the

maximum kinetic energy to be

less than 75 ft-lbf.

Analysis

3.4 The recovery system electrical circuits will be

completely independent of any payload

electrical Circuits.

The recovery system electrical

circuits will be completely

independent of any payload

electrical circuits.

Demonstration

3.5 All recovery electronics will be powered by

commercially available batteries. Commercially available 9 volt

batteries will power all recovery

electronics.

Visual

3.6 The recovery system will contain redundant,

commercially available altimeters. The term

“altimeters” includes both simple altimeters and

more sophisticated flight computers.

The launch vehicle will contain

redundant PerfectFlite

Stratologger altimeters.

Visual

3.7 Motor ejection is not a permissible form of Motor ejection will not be used N/A


primary or secondary deployment. as a method of deployment.

3.8 Removable shear pins will be used for both the

main parachute compartment and the drogue

parachute compartment.

Removable shear pins will be

used for all parachute

compartments.

Visual

3.9 Recovery area will be limited to a 2500 ft.

radius from the launch pads. Estimated drift distance will be

analysed and limited to 2500 ft

or less.

Analysis

3.10 An electronic tracking device will be installed

in the launch vehicle and will transmit the

position of the tethered vehicle or any

independent section to a ground receiver.

3.10.1. Any rocket section, or payload

component, which lands untethered to the

launch vehicle, will also carry an active

electronic tracking device.

An Xbee will be used to track

all separate components of the

launch vehicle and payload.

Demonstration

3.10.2. The electronic tracking device will be

fully functional during the official flight on

launch day.

The electronic tracking device

will be tested prior to the

official flight on launch day

Demonstration

3.11 The recovery system electronics will not be

adversely affected by any other on-board

electronic devices during flight (from launch

until landing).

3.11.1. The recovery system altimeters will be

physically located in a separate compartment

within the vehicle from any other radio

frequency transmitting device and/or magnetic

wave producing device.

The altimeters will be located in

a separate compartment from

RF transmitting or Magnetic

wave producing devices.

Demonstration

3.11.2. The recovery system electronics will be

shielded from all onboard transmitting devices,

to avoid inadvertent excitation of the recovery

system electronics.

Avionics bays will be shielded

from transmitters onboard the

payload in order to prevent

interference.

Demonstration


shielded from all onboard devices which may

generate magnetic waves (such as generators,

solenoid valves, and Tesla coils) to avoid

inadvertent excitation of the recovery system.

The electronics will be shielded

from solenoids in the ejection

methods and other systems in

the rocket.

Demonstration


shielded from any other onboard devices which

may adversely affect the proper operation of the

recovery system electronics.

A Faraday cage will be used for

the avionics bay at the interface

of the payload and booster

sections to shield recovery

electronics from electronics on

the rover.

Demonstration

4.1 Each team will choose one design experiment

option from the following list. Deployable rover chosen. Verified

4.2 Additional experiments (limit of 1) are allowed,

and may be flown, but they will not contribute

to scoring.

No additional experiments. N/A

4.3 If the team chooses to fly additional

experiments, they will provide the appropriate

documentation in all design reports, so

No additional experiments. N/A


experiments may be reviewed for flight safety.

Option 1: Target detection; Option 2:

Deployable rover; Option 3: Landing

coordinates via triangulation.

4.5 Deployable rover

4.5.1. Teams will design a custom rover that

will deploy from the internal structure of the

launch vehicle.

A pneumatic piston or similar

piston assembly will be used to

deploy a capsule containing the

rover

4.5.2. At landing, the team will remotely

activate a trigger to deploy the rover from the

rocket.

An XBee will be used to

communicate between a ground

station and the rover. A

command will be sent to actuate

the ejection method. The

ground station will be capable

of verifying if information was

received and actuation occurred.

Ground station, visual

4.5.3. After deployment, the rover will

autonomously move at least 5 ft. (in any

direction) from the launch vehicle.

Servos will be used to turn

wheels for locomotion. Rotational encoder,

visual measurement

4.5.4. Once the rover has reached its final

destination, it will deploy a set of foldable solar

cell panels.

Solar panels will rotate along an

axis perpendicular to their face,

thus unfolding in the same

plane.

Visual

5.1 Each team will use a launch and safety

checklist. The final checklists will be included

in the FRR report and used during the Launch

Readiness Review (LRR) and any launch day

operations.

A preliminary checklist will be

developed over the course of

the fall semester by the

leadership. Subgroup leads will

ensure that their subsystems are

accounted for, and safety officer

and team lead will provide

double checks as well as safety

and general checklists. A final

checklist will be completed

before FRR

Demonstrated at

milestone reviews

5.2 Each team must identify a student safety officer

who will be responsible for all items in section

5.3.

The team’s safety officer will

be Rebecca Holleb Verified

5.3 The role and responsibilities of each safety

officer will include, but not limited to:

5.3.1. Monitor team activities with an emphasis

on Safety during:

5.3.1.1. Design of vehicle and payload Rebecca will be present during

all team meetings, including

vehicle and payload design

meetings.

N/A

5.3.1.2. Construction of vehicle and payload Rebecca will be present during

fabrication of vehicle and

payload during fabrication

meetings.

N/A


5.3.1.3. Assembly of vehicle and payload Rebecca will be present during

assembly of vehicle and

payload during assembly

meetings.

N/A

5.3.1.4. Ground testing of vehicle and payload Rebecca will be present for all

ground tests. N/A

5.3.1.5. Sub-scale launch test(s) Rebecca will be present for all

sub scale launches. N/A

5.3.1.6. Full-scale launch test(s) Rebecca will be present for all

full scale launches. N/A

5.3.1.7. Launch day Rebecca will be present at

launch day. N/A

5.3.1.8. Recovery activities Rebecca will be present for

recovery activities. N/A

5.3.1.9. Educational Engagement Activities Rebecca will be present for

educational engagement

activities.

N/A

5.3.2. Implement procedures developed by the

team for construction, assembly, launch, and

recovery activities

Rebecca has assisted in the

formulation of these procedures

and will ensure they are

implemented.

Demonstration at ground

tests, launches and

internal performance

evaluations

5.3.3. Manage and maintain current revisions of

the team’s hazard analyses, failure modes

analyses, procedures, and MSDS/chemical

inventory data

Rebecca will work closely with

the club safety officers for

guidance with these matters

towards the end of properly

maintaining understanding of

the current hazards and failure

modes.

Demonstration in

milestone reviews and

internal reflections on

performance

5.3.4. Assist in the writing and development of

the team’s hazard analyses, failure modes

analyses, and procedures.

Rebecca contributes a great deal

to these sections. Verified

5.4 During test flights, teams will abide by the rules

and guidance of the local rocketry club’s RSO.

The allowance of certain vehicle configurations

and/or payloads at the NASA Student Launch

Initiative does not give explicit or implicit

authority for teams to fly those certain vehicle

configurations and/or payloads at other club

launches. Teams should communicate their

intentions to the local club’s President or

Prefect and RSO before attending any NAR or

TRA launch.

AIAA at Northeastern

University remains in good

communication with Central

Massachusetts Spacemodeling

Society (CMASS), NAR

chapter 464. We attend many

launches per season and have

good relationships with the

RSO.

Demonstration of

requirement verification

at test launches

5.5 Teams will abide by all rules set forth by the

FAA. Safety officer will be familiar

with FAA rules. Team will refer

to the club design safety

committee for FAA guidelines

and compliance

Review meeting with

design safety committee


Table 5.2.2 : Verification Plan Cont.

Requirement Justification Completion Plan Verification

Vehicle

Payload bay open to air and

unencumbered at landing Required for proper

payload deployment Payload section will be connected at

one end to recovery system while at

the other a charge will separate the

two sections, leaving the payload exit

open to the air.

Demonstration at

ground tests, subscale,

and full scale

launches

Payload undamaged by

deployment method Required for robust

payload deployment and

for reusability

Designs are constrained by ability to

protect payload consistently from

damage.

Simulation,

demonstration at


and full scale

launches

Payload deployment and

vehicle recovery system

supported by redundant

systems

Encouraged by internal

club policies and club

design safety committee

Altimeters will have redundant units;

ejection method will have backup and

fail-safe systems.

N/A

Ejection System and Payload Experiment

Payload ejection system is

robust and reliable so that

successful deployment occurs

every time

Required for reusability,

required for success at

competition

Ejection method will not be chaotic,

controlled designs only will be

considered.

Demonstration at


and full scale

launches

Ejection system and payload

shall not weigh more than 3

kg combined

Required for launch

vehicle design to ensure

apogee of 5280 ft

Simulation,

measurements

Proper safety precaution will

be taken when handling

pressurized gas

Specific to our

experiment and required

by handbook

Team members working with

pressurized gas will become familiar

with all guidelines and consult club

safety officers for guidance.

N/A

Payload capsule will fully

open upon deployment from

airframe

Required for safe

deployment of rover

without damage,

considered success

criteria of experiment

Systems of springs and latches will be

used to open the capsule and ensure it

stays open and is open enough for

payload clearance.

Demonstration at


and full scale

launches

Payload will deploy its

counter-torque support wheel

upon exiting the airframe

Considered success

criteria of experiment Support will be controlled by servo

which will be directed to deploy

support once exit is confirmed.

Demonstration at


and full scale

launches

Rover will be capable of

moving 5 ft away from rocket

and overcoming terrain

Internally derived

requirement to amend

handbook requirement

based on knowledge of

launch site terrain

Wheels will be designed and tested on

loose surface material. Demonstration at


and full scale

launches


Rover will be capable of

deploying solar panels in

correct orientation

Considered success

criteria of experiment Accelerometer will detect orientation

and instruct solar panels on correct

deployment.

Demonstration at


and full scale

launches

Recovery

The payload section will not

fall at more than 18 ft/s

Specific to protection of

payload Recovery system will be designed

with payload weight in mind such that

the maximum terminal velocity is not

exceeded.

Simulation,

demonstration at


and full scale

launches

Safety

Club safety officers will be

present during lab time and

fabrication

Required club, lab and

university policy Fabrication meetings will only occur

during lab hours. Internal check-ins

with club safety

officers

Club members will attend

safety presentation given by

club safety officers

Required by club policy Safety presentations given at

beginning of semester. Verified

Club design safety committee

will review design and final

craft before launch for

adherence to safety and NAR

and FAA requirements

Required club policy Internal design plan review and safety

plan already completed, flight

readiness reviews of any launch or

ground test will be given by design

safety committee.

Internal check-in with

design safety

committee

General

Upperclassmen will

encourage engagement in

freshmen and sophomores

Required to maintain high

retention rates among

underclassmen and for

succession planning

Leadership will be friendly and make

effort to befriend younger members,

snacks provided at certain meetings.

Supportive language and constructive

advice will be given in engineering

learning settings

Analysis of retention

data and from verbal

feedback from

freshmen

Team leadership will meet

weekly to discuss progress

and update one another on

changes made to designs

Required for successful

execution and

administration of project

Leadership meetings occur at 7:00 PM

on Thursdays for the fall semester. N/A


5.3. Budgeting and Timeline Line Item Budget

Our line item budget for the spring semester is given below. Funding Plan

AIAA at NU has submitted requests for funding from multiple sources for the 2018 NASA University Student Launch Initiative. To cover material goods that will be used in the

development and construction of our full-scale rocket and payload, we have submitted a budget to

Northeastern’s Student Government Association (SGA). This money will become available to us in late

January, at which point we will purchase our materials. Until then, we have set aside provisional funds

from last semester to purchase high-priority materials such as fiberglass for body tube. For travel

expenses and STEM outreach related costs, we have applied for a grant of $3,000 from Northeastern’s

COE Scranton fund. For an additional source of funding, we will also apply to conduct a campaign on

Northeastern’s HuskyStarter crowdsourcing platform. The money from the Scranton Fund and

HuskyStarter will go towards transportation fees. We will present our project to SGA’s finance board and

follow their specified protocol in order to request funding for our budget. We have high confidence that

our request of $3,486 will be granted. Our next step for funding is applying to Northeastern’s

HuskyStarter. This crowdsourcing platform allows projects to be exposed to a community of thousands of

potential donors, as well as generates interest within the community. We will have to complete an online

application, after which, we will present a photo or video presentation to potential donors and document

the progress of our launch vehicle. We have been successful with HuskyStarter in the past, so we believe

we will get this funding as well. These sources will fully support our team’s financial needs through

completion of the project. Our expected travel expenses are listed below.

Table 5.3.1 : Travel Expenses

Item Vendor Price Quantity Cost

January 14th MDRA

Gas (1 car) Various $70 1 Car $70

Tolls (1 car) Various $52 1 Car $52

Hotel (1 room) Courtyard Marriott $100 1 Night $100

February 17th-18th MDRA

Gas (2 cars) Various $70 2 $140

Tolls (2 cars) Various $52 2 $104

Hotel (2 rooms) TBD $200 1 Night $200

Huntsville


4 Hotel Rooms Embassy Suites $400 5 Nights $2000

Shenandoah Campsite Nat’l Parks $15 3 Sites $45

Great Smoky Mts. Campsite Nat’l Parks $35 1 Site $35

Gas (3 cars) Various $160 3 Cars $480

Tolls (3 cars) Various $60 3 Cars $180

Total $3316

Table 5.3.2: Launch Vehicle Budget


Table 5.3.3 : Payload Budget

Table 5.3.4 : Electronics Budget and Total

Table 5.3.5 : Project Schedule

Task Start Date Duration End Date

Winter Break 12/16/2017 23 1/8/2018

Fabricate Final Rover 1/9/2018 39 2/17/2018

CDR 1/5/2018 7 1/12/2018

*CDR Telecon 1/17/2018 1 1/18/2018

Ground Test Full Scale LV 2/12/2018 5 2/17/2018

Full Scale Test Launch 2/17/2018 1 2/18/2018

Test Launch Analysis 2/19/2018 2 2/21/2018

Update LV Design (if needed) 2/19/2018 4 2/23/2018


Update Rover Design (if needed) 2/19/2018 4 2/23/2018

Implement LV Changes 2/23/2018 10 3/5/2018

Implement Rover Changes 2/23/2018 10 3/5/2018

FRR 2/26/2018 7 3/5/2018

*FRR Telecon 3/7/2018 1 3/8/2018

*Ground Test (Tentative) 3/5/2018 5 3/10/2018

*Test Launch (Tentative) 3/10/2018 1 3/11/2018

*Test Launch Analysis (Tentative) 3/12/2018 2 3/14/2018

Update and Implement LV Designs 3/14/2018 19 4/2/2018

Update and Implement Rover Designs 3/14/2018 19 4/2/2018

Logistics Review/Finalization 3/26/2018 4 3/30/2018

Travel to Competition 4/2/2018 2 4/4/2018

Competition Week 4/5/2018 3 4/8/2018

Figure 5.3.6. : Project Schedule Gantt Chart

Date post:	20-Apr-2020
Category:	Documents
Upload:	others
View:	3 times
Download:	0 times

University Student Launch Initiative 2017-2018 Critical ... · Justification for material...

Documents